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Apollo Extension Systems—Lunar Excursion Module

Phase B Final Report

N76-70962[(NASA-CR- 0879) APOLLO EXTENSION SYSTEMS:LUNAR EXCURSION MODULE, PHASE B. VOLUME 16-

AND MISSION OPEEATIONS FinalUo£°rt (Grumaan Aircraft Engineering Corp.) Unclas

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Vol. XVI Prelaunch & Mission Operations flJ)

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Copy No.11

Apollo Extension Systems — Lunar Excursion ModulePhase B Final Report

to

National Aeronautics and Space AdministrationManned Spacecraft Center

Advanced Spacecraft Technology DivisionHouston, Texas 77058

byGrumman Aircraft Engineering Corporation

Bethpage, New York

This document contains information affecting the national defenseUnited States within the meaning'ohttie Espionage Laws, TitleSections 793 and 794, the transmissionxlKre^elation of whichjfljemy mannerto an unauthorized person is prohibited by}

Vol. XVI Prelaunch & Mission Operations

Contract No. NAS 9-4983ASR 3788

8 December 1965

Preface

This report presents the results of the Phase"B" Preliminary Definition Study (ContractNAS 9-4983) of the Lunar Excursion Module(LEM) and its modifications and additions, asnecessary, for use in the Apollo ExtensionSystems (AES). This use includes a Laboratoryfor Earth and lunar orbital missions, and aShelter, a Taxi and a Truck for extended-staylunar surface missions. The overall objective ofthis study was to conduct sufficient analysesto provide a basis for selection by NASA of asingle concept for each mission for final defini-tion and development.

The study results are distributed in the vol-umes listed below in the following manner:Volume I contains a summary of the Prelimin-ary Project Development Plan (PDP) withemphasis on estimates of the program costsand schedules. This volume was submitted on30 October 1965, one month in advance of theremaining final documentation. Volume II is abrief summary of the overall study. VolumesIII through XVI contain the design analyses,

-preliminary specifications, and operations an-alyses for each of the AES/LEM vehicle types.Volumes XVII through XXVI contain prelim-inary project planning data in the areas ofmanagement, manufacturing, development test-ing, and support.

It was necessary to base the preliminaryproject planning data, including estimatedcosts, on a single configuration for each of theAES/LEM vehicle types. Since these PDP datawere required by the end of October, the con-figurations had to be selected at the mid-pointof the study, before the configuration studieshad been completed. These configurations havebeen called "baseline" configurations. The con-tinuing design analyses in the second half ofthe study have resulted in recommendedchanges to the baseline configurations. VolumesIII through VI describe the "recommended"configurations, the baseline configurations, andsome additional alternates which were studied.It is anticipated that NASA will make a selec-tion from these configurations, and that theseselections will then be the new baseline con-figurations for the next phase of AES definitionstudies.

The scope of this study included integrationof the experimental payloads with the Shelterand Taxi, but did not include study of the inte-

gration on individual LEM Laboratory flights.At approximately the mid-point of the study, anaddendum was written with the objective ofproviding support to the NASA Mission Plan-ning Task Force for study of the Phase I Lab-oratory flights. The schedule for the addendumcalls for completion of these mission planningstudies in January, 1966. Therefore, the adden-dum efforts are not described in this report.

The volumes which comprise this report areas follows:

/ Phase B Preliminary Definition Plan(30 Oct 1965)

II Preliminary Definition StudiesSummary

III Phase I LaboratoryDesign Analysis Summary

IV Phase II LaboratoryDesign Analysis Summary

V Shelter Design Analysis SummaryVI Taxi Design Analysis Summary-

VII Truck Design Analysis SummaryVIII Phase I Laboratory

Master End Item SpecificationIX Phase II Laboratory

Master End Item SpecificationX Shelter Master End Item Specification

XI Taxi Master End Item SpecificationXII Phase I Laboratory Experimental

Payload Performance &Interface Specification

XIII Phase II Laboratory ExperimentalPayload Performance &Interface Specification

XIV Shelter Experimental PayloadPerformance & Interface Specification

XV Taxi Experimental PayloadPerformance & Interface Specification

XVI Prelaunch & Mission OperationsXVII Manufacturing Plan

XVIII AES Modifications to LEMQuality Control Program Plan "

XIX Ground Development Test PlanXX Support Equipment Specification

XXI Facilities PlanXXII Support Plan

XXIII Transportation PlanXXIV Training Equipment RequirementsXXV Support Equipment Requirements

XXVI Management Plan

(XVI) 111

. CONTENTS -

Section ' • . . • ' . • . .Page

1 . INTRODUCTION

'. . 1.1 • BACKGROUND 1-11.2 STUDY OBJECTIVES . . . . . . . . . . . . . . . 1-11.3 . ' STUDY APPROACH . . . . . . . . . . . . . . . . . 1-1

2 MISSION OPERATIONS . ' . ' • , • ' .

2.1 FLIGHT OPERATIONS LOGIC • ... . .... . 2-12.1.1 . General • • • . ' • • > • • ........ 2-12.1.2 Test Requirements . . . . . . . . . . . . . . . 2-12.1.3 Text Logic . .................. 2-2

; . . 2.2 PHASE I LAB MISSIONS... . . . . . . . . . . . . 2-22.2.1 Earth Orbit, Saturn V Launch .'. ....... 2-2

.• 2.2.2 Earth Orbit Rendezvous, Saturn IB Launch . ... 2-5..2.2.3 Earth Polar Orbit, Saturn V Launch .. 2-8

2.2.4 . Earth Synchronous Orbit, Saturn V .Launch ... 2-102.2.5 Lunar Orbit, Saturn V Launch. . . . . . . . . . 2-13

•'2.3. ' PHASE II LAB MISSIONS . . . . . . . . . . . . . 2-2*1" 2.3>1 Earth Orbit, Low-Inclination, Saturn V Launch'. 2-24

2.3.2 Earth Orbit Rendezvous, Saturn IB Launch . . . 2-282,3.3- Earth Polar Orbit, Saturn V Launch ...... 2-302.3.4 Earth Synchronous Orbit, Saturn V Launch . .'... 2-322.3.5 .High-Inclination Lunar Orbit, Saturn V Launch . 2-342.4 '". SHELTER MISSIONS . . . . . . . . . .. . . . . . 2-372.5 . TAXI MISSIONS 2-^22.6 TRACKING REQUIREMENTS . . ' 2-452.6.1. Ground Rules ; . . . . . . . . 2-^52.6.2 Assumptions 2-U52.6.3 . Summary of MSFN Capabilities 2-^52,6.U Earth Orbital Lab Missions 2-462.6.5. . Ground Tracking Analysis for Lunar Missions . . 2-472.6.6 . Lunar Mission Communications Operations .... 2-492.7 OPERATIONAL DATA HANDLING 2-832.7.1 Assumptions .................. 2-83

. 2.7.2 Basic Spacecraft Data Requirements ...... 2-832.7.3 ' Experiment Data Interface . . . 2-832.7.4: Operational Data Flow . . . . 2-83

. 2.7-5 Instrumentation Subsystem , 2-83

3 PRELAUNCH OPERATIONS.

3.1 CHECKOUT RATIONALE .'...' 3-13.1.1 • Concepts and Criteria . 3-13.2 SUBSYSTEM CHECKOUT APPROACH 3-4

(XVI) iv

CONTENTS

Section ' Page

3.2.1 Propulsion/RCS Subsystems 3-43.2.2 Environmental Control Subsystems 3-53.2.3 Guidance Navigation & Control Subsystem 3-73.2.U Electrical Power Subsystem . .'.'.'. . 3-103.2.5 Communications Subsystem, . 3-113-2.6 Instrumentation Subsystem 3-123.2.7 Structural/Mechanical & Electro-Explosive Devices

Subsystems . " . . . . 3-133.2.8 Crew Provisions Subsystems. 3-1^3.2.9 Experiment Interfaces 3-153-3 CHECKOUT OPERATIONS . 3-283.3-i Assumptions . . " . . . . . 3-283.3-2 ' Phase I Lab . . ..... ... . ". . . 3-293.3.3 Phase II"Lab . . . . . . . . . . . . . . . . . . 3-313-3.4 Shelter . . . . . . . . 3-3U3.3.5 Taxi. . . . / . . . . . . . . . . . . . 3-363*U ' PRELAUNCH OPERATIONS SCHEDULE ....'. 3-hi3.4.1 Assumptions . . . :, "" 3-4l3-5 CHECKOUT REQUIREMENTS AT .KSC ' . . ..'.'. 3-1*53.5.1 Facility Description ...... . . . , . 3-453.5.2 Facility Requirements-. . .- ".'.. 3-1*73.5«3 Acceptance Checkout Equipment (ACE) Requirements. 3-503.5.^- Ground Support Equipment (GSE). Requirements . . . 3-51

1* REFERENCES . . . . ... . .,..". . . .' ."" 4-1

5 ABBREVIATIONS .AND SYMBOLS . ..'. . . . . ..: ....'.' 5-1

(XVI)

Tables

2.1-1 AES S/C Test Requirements Summary 2-l62.1-2 AES Ground Development Testing 2-182.1-3 AES Ground Test Article Usage for Flight Support . 2-202.1-U AES Flight Mission Prerequisites . .' 2-222.6-1 Unified S-Band Station Capabilities 2-532.6-2 Summary of Ground Station Coverage - Polar Orbit and -83° Orbit . 2-552.6-3 Summary of Ground Station Coverage - Lunar Orbit 2-572.6-H Possible Ground Station Modes for Shelter/Taxi Lunar Stay .... 2-582.7-1 Real Time Status Data Handling Requirements „ . 2-862«,7-2 Vehicle Total Measurement Points. .2-932»7-3 Telemetered Measurements. .... 2-952.7-U GSE - Vehicle Access Measurements 2-972.7-5 Vehicle Flight Measurements . . . . 2-99

3.2-1 Propulsion/RCS Checkout Configurations 3-l63.2-2 ECS Checkout Configurations 3-173.2-3 FCS Checkout Configurations 3-l83.2-1+ EPS Checkout Configurations 3-193.2-5 Communications Checkout Configurations 3-203.2-6 Instrumentation Checkout Configurations 3-213»2-7 Structural/Mechanical & EED Checkout Configurations 3-22.3 ,,2-8 Crev Provisions Checkout Configurations 3-233.2-9a Experiment Interface Assumptions & Categories „ . . 3-23<,2-9b Experiment Interface Checkout.Objectives & Categories 3-253.2-9c Experiment Approach & Categories . 3-26

(XVI) vi

Illustrations

2.1-1 AES Flight Mission & Constraint Logic . . 2-232.6-1 Lunar Planar Projection . . . . . . '2-592.6-2 .Ground Track - 28.5° Incl. Orbit. ;..... .. .......... 2-6l2.6-3 Ground Station Visibility - 28.5° Incl. Orbit . . . . 2-632.6-4 Ground Track - 90° Incl. Orbit 2-692.6-5 Ground Track - 83° Incl. Orbit 2-712.6-6 Ground Track Synchronous Orbit. ....'... 2-732.6-7 Ground Station Visibility - Synchronous Orbit 2-752.6-8 Ground Station Visibility - Lunar Orbit ..... 2-772.6-9 Ground Station Visibility - Lunar Polar Orbit .......... 2-792.6-10 Effect of Lunar Position on Earth Coverage 2-8l2.6-11 Shelter/Taxi Operations in Lunar"Orbit. . , 2-82

3.3-1 KSC Faciiity/AES Vehicle Flows 3-393.4-1 Prelaunch Operations Schedule3.5-1 KSC Facility Loading Chart. ................... 3;433.5-2 Vehicle/Facility ACE Utilization. . . . . . . . . . . . . / . . . '"3-553.5-3 KSC ACE Utilization & Prelaunch Operations. . ......... .3-573.5-4 Vehicle/Facility/ACE Loading Summary. . . ..... ."'3-593.5 AES-Lem Checkout Equipment Requirements ........ 3~6l

/uimman.

(XVI)1-1

1. INTRODUCTION

1.1 BACKGROUND

The Prelaunch and Mission Operation Report presents the results of a preliminaryanalysis of the operations requirements for the AES LEM Program. The study hasbeen carried out in response to the Statement of Work for Contract NAS 9- 983. Thetask has been divided into two concurrent study efforts: one concerned withmission and mission related operations such as tracking requirements; and the otherwith prelaunch and launch operations.

1.2 STUDY OBJECTIVES

The Grumman study objective for operations planning was to conduct sufficient anal-ysis to provide NASA a basis for making a sound judgment on follow-on operationsactivities. The purpose of this study was not to concentrate on the optimization ofone particular solution, but rather to explore a number of possible approaches andselect the most promising. The primary effort has been directed toward establishingrealistic criteria, identifying problem areas and constraints, and carrying theplanning to the point where solutions appear to be practical. AES-LEM spacecraftoperational requirements will be a foundation for developing detailed operationalrequirements as spacecraft and experiment configurations are made more definitive.

1.3 STUDY APPROACH

During the 5-month Phase B study, the Baseline Spacecraft Configurations and theRecommended Spacecraft Configurations were considered in operations planning.These configurations are identified in the-"Vehicle Design Analysis Summaries"(Phase I Lab, Vol. III;. Phase II Lab, Vol. IV; Shelter, Vol. V; Taxi, Vol. VI).Volume I, "Phase B Preliminary Definition Plan" dated 29 October 1965, reflects theoperations planning for the Baseline Configurations. Subsequently, all planningefforts were reviewed and updated in accordance with the Recommended Configurations.•All prelaunch and mission operations planning presented in this volume reflects theRecommended Configurations.

The ground rules, guidelines, and assumptions pertinent to prelaunch and missionoperations are included in proper context within the appropriate sections of thisvolume.

Section 2 of this report presents the preliminary flight mission planning informa-tion. First, the flight operations logic including flight test requirements andprerequisite ground tests are summarized. Next, the Phase I Lab, Phase II Lab,Shelter, and Taxi missions are categorized and described. Emphasis is given tospacecraft test objectives and rationale. After each mission category is presented,the mission related tracking, operational data handling and instrumentation re-quirements are discussed.

Section 3 of the report presents the preliminary prelaunch operations planning.The checkout rationale pertinent to the development of prelaunch checkout plans,sequences, and vehicle flows for the AES-LEM spacecraft is summarized. After

(XVI)1-2

describing the checkout approach for each AES-LEM subsystem, the checkout flow foreach AES-LEM spacecraft-launch vehicle combination is developed. These individualPhase I Lab, Phase II Lab, Shelter, and'Taxi checkout flows are then integrated intoan overall KSC launch operations schedule. The checkout operations are analyzed toassess the KSC requirements for the AES-LEM Program with respect to those of theApollo Program. The checkout requirements for the overall AES-LEM Program are iden-tified for KSC facilities, ACE, and ground support equipment. The Phase I Laboperations planning and scheduling, presented in Section 3; emphasizes the AES-LEMspacecraft requirements and assumes a nominal experiment complement. During- thecourse of the Addendum study, the effects of .Lab experiments defined by NASA foreach Phase I Lab will be investigated.

(XVI)2-1

2. MISSION OPERATIONS

2.1 FLIGHT OPERATIONS LOGIC

2.1.1 General

This section presents the operational aspects of mission planning. The flighttest requirements summarized herein together with the scientific mission objec-tives will describe the total mission requirements for the AES-LEM Program.Mission planning studies emphasizing mission related requirements on spacecraftdesign and performance capabilities are presented in Vol. Ill through VII.

2.1.2 Test Requirements

The prime purposes of the AES missions are to conduct extensive experiments forbasic scientific research, develop and qualify new systems and subsystems, showspacecraft and crew capability to withstand long-duration missions, and developspace operating procedures and techniques. In lieu of performing special flightdevelopment missions, the AES Program will utilize the prerequiste Apollo Programflight development, in conjunction with a comprehensive AES Ground DevelopmentTest Program, to relieve AES test constraints. However, flight test objectiveswill be integrated within the flight operations of the initial flight of eachAES-LEM/mission combination. The chief requirements for this are:

• To demonstrate system performance for environmental conditions notattainable to a satisfactory degree on the ground.

• To assure subsequent mission success of each similar AES-LEM.• To assure maintaining the high density launch schedule.

Table 2.1-1 indicates the flight test requirements associated with each of theAES-LEM. These requirements are further defined in terms of test objectives foreach type of AES-LEM/mission combination in Paragraphs 2.2 through 2.5. As thescientific mission objectives are defined in detail, the flight test requirementswill be reviewed and considered in the light of total flight mission requirements.

Since each type of AES-LEM is scheduled for a number of flight missions, it isdesirable to satisfy the spacecraft test objectives as early in the flight programas practical, e.g. on the first flight of each vehicle type. The planned numberof AES flight missions are:

AES-LEM LAUNCH VEHICLESaturn IB Saturn V

Phase I Lab 2 3Phase II Lab k 7Shelter - .3Taxi - 3

(XVI)2-2

2.1.3 Test Logic

The AES flight mission and constraint logic is presented in Fig. 2.1-1, whichshows the dependence on Apollo flights and the AES Ground Development Test Programs.The AES mission prerequisite testing is a combination of the AES ground develop-ment tests which are detailed in Vol XIX and summarized in Tables 2.1-2 and 2.1-3,and the basic Apollo Program test prerequisites listed in Table 2.1-A.

2.2 PHASE I LAB MISSIONS

2.2.1 Earth Orbit, Saturn V Launch

2.2.1.1 Mission Description

The generalized Phase I Lab flight profile is an earth orbital mission at analtitude of 200 n.mi, and orbital inclination of 28.5 or 50 deg, and a flightduration of ik days. A Saturn V launch vehicle is used to boost the spacecraftinto the 200-n.mi circular orbit for experiment operations. A three-man crewis required for these missions.

The prime purpose of the Phase I Lab missions is to conduct extensive experimentsfor basic scientific research, develop and qualify new systems and subsystems, andqualify equipment for subsequent missions. These-missions are also intended toshow spacecraft and crew capability to withstand long-duration space flights, anddevelop space operating procedures a n d techniques. • ' • • • . . - . - .

2.2.1.2 Spacecraft Test Objectives

The following spacecraft•test objectives have been derived from the logic ofParagraph 2.1 and are a general group that pertain to all Phase I Lab missions.

1.

OBJECTIVE

Demonstrate the performance ofthe modified LEM structure andinsulation over a 1^-day Earthorbital Lab mission.

2. Demonstrate the pointingcapability of Lab FlightControl System (FCS).

2.

RATIONALE

Previous Apollo LEM test,flights will haveshown that the basic LEM structure andinsulation are capable of withstanding theboost environment plus short duration staysin a thermal vacuum environment. Modi-fications to the structure caused by equip-ment additions and deletions plus - additionsof thermal insulation require demonstrationof the new configuration's ability to •support a 14-day mission in the near spaceenvironment.

The Apollo FCS will have been tested overthe complete range of LEM inertias. How-ever, the LEM will not have been requiredto maintain the stability that variousEarth oriented applications require. Theability of the modified EGA and ATCA tomaintain a 0.1-deg, one-pulse, limit cyclemust be demonstrated. In addition, the

OBJECTIVE

2. (Cont'd)

Demonstrate operation of themodified LEM subsystems afterbeing subjected to boostenvironment.

Demonstrate the capabilityof the modified LEM ECS andcrew provisions to supportlife over a lU-day Earthorbital Lab mission.

Evaluate the Lab data manage-ment systems including instru-mentation, communications, andMSFN/MCC performance in han-dling experimental data.

6. Evaluate extravehicular crewoperations.

RATIONALE

modified Lab systems, plus the requirementof the Lab to maintain Lab/CSM stability,require the system to be completelychecked out in flight.

3, The basic Apollo LEM Subsystem configur-ation will have been flown on both SaturnIB and Saturn V boosters. Structure andsupport modifications require that theloads and vibrations imparted to the Labby the Launch Vehicle and by the Lab tothe various equipment areas be investi-gated. ECS and EPS Subsystem additions,plus deletion of the Propulsion Sub-systems, require that the boost environ-ment effects on these modified subsystemsbe demonstrated.

4. The Lab ECS subsystem is a modifiedversion of the LEM ECS. The modificationsare chiefly in the form of added comsum-ables and tankage. The ECS must be shownto have the capability of life support fora lU-day mission, and consumable usage fora il -day mission must be established. Itmust also be shown that the ECS can pro-vide atmosphere circulation for the Lab/CSM combination. The ability of the ECSto provide equipment and vehicle thermalcontrol during long-duration flights inconjunction with maintaining the crew mustbe investigated.

5- Although the Instrumentation and Communi-cations Subsystems will have been checkedout previously during Apollo Earth orbitaland lunar missions, the mass of experimentdata that must be handled will not havebeen approached. Methods of storing,transmitting, and ascertaining properground acquisition of large quantities ofdata must be established for periods%ofminimum ground coverage.

The ability of the crew to perform usefulextravehicular activities will not havebeen evaluated previously to the extentrequired in the AES missions. The abilityto transfer equipment from and to the Lab,perform repairs, and carry out experimentsand potential rescue operations must be

(XVI) 2-1*

OBJECTIVE

6. (Cont'd)

RATIONALE

confirmed as early as practical in theAES Program. Changes in the internaland external Lab configuration requiredemonstration of crew ingress, egress,and extravehicular operations.

2.2.1.3 Earth Orbit, Saturn V Launch, Mission Profile (200 N. Mi)

MISSION PHASE NOMINAL PHASE TIME(Hr: Min: Sec)

00:12:00Saturn V Ascent to Orbita) S^IC firingb) S-II Firingc) LES Jettispnd) S-IVB firing

S-IVB Stabilization . 03:00:00a) CSM C/0

Transposition & Docking 00:30:00a) CSM-S-IVB separationb) Transpositionc) CSM-LEM Lab dockingd)•• LEM-Lab extraction from

S-IVB

LEM Lab C/0 01:30:00a) Crew transfer to Labb) Lab C/0

Earth Orbit Lab Experiment 336:00:00Operations . (lU days)

OPERATIONAL TEST OBJECTIVE

None

None

None

Demonstrate operation ofmodified LEM subsystemsafter being .subjected toboost environment.

Demonstrate the:• Pointing- capability of

FCS. -:'- '" -•'• Modified LEM ECS-and Crew

Provisions capability tosupport life for Ik-daymission.

• Performance of modifiedLEM structure and insula-tion over l4-day mission.

Evaluate the:• Lab data management sys-

tem.• Extravehicular crew

operations.

(XVI)2-5

MISSION PHASE NOMINAL PHASE TIME OPERATIONAL TEST OBJECTIVE(hr: Min: Sec)

6. Lab Shutdown & Equipment 01:30:00and Data Transfer to CSM

7- Preparation for CM Entry 00:10:00a) Lab jettisonb) SPS deorbit maneuverc) SM jettison

8. Entry 00:11:00

9. Parachute descent 00:07:00

10. Post landing 'through VariableS/C retrieval

2.2.2 Earth Orbit Rendezvous, Saturn IB Launch


Saturn IB launched Labs will be part of dual launch Earth orbital rendezvous missions,The Lab will be launched into a 200-n.mi Earth orbit at an inclination of 28.5 deg.The launch vehicle configuration will be that of the Apollo 206A/LEM-1 mission,(Ref. U.I), and consists of Lab, boilerplate CSM, SLA, and LES. The CSM will belaunched with the crew and will rendezvous and dock to the LAB. Manned missionduration is lU days and a three-man crew is required. The primary purpose of thistype mission is to extend the capability of the CSM and provide a greater orbitalpayload than that available with a single Saturn IB launch.


In addition to the basic objectives in Paragraph 2.2.1, the following spacecrafttest objectives are associated with an Earth orbital rendezvous mission profile.

OBJECTIVE RATIONALE

1. Demonstrate integrated mission 1. The dual launch concept requires coordi-control from countdown through nated efforts during the prelaunch check-dual launching, equal period out to assure proper timing relationshipsorbit capability, rendezvous, between the Lab and CSM launches. Theand docking. two pad (3U & 37B) launch operations, as

well as the coordinated tracking operationsfor two separate vehicles, will have notbeen demonstrated during the basic ApolloProgram.

j/utmman.

(XVl)2-6

2.

OBJECTIVE

Demonstrate the ability ofthe Lab ECS, EPS, Instru-mentation and Communicationsequipment to perform asrequired during unmannedorbital storage period.

RATIONALE

2. The ability of the Lab ECS to maintainthermal control 'during the unmannedstorage period, and of the EPS to supplysufficient power for the storage periodmust be shown. The Instrumentation andCommunication Subsystems' ability tooperate after prolonged exposure to athermal vacuum environment must be shown.Ground tests can give an indication of thesubsystem requirements for prolongedstorage; however, only during space opera-tions can the effects of storage in athermal vacuum environment- be. demonstrated.

3- Long-duration orbital storage may have adetrimental effect on the Lab FCS, especi-ally if the RCS is pressurized and is usedfor extracting the Lab from the S-IVB/SLAprior to storage. The effects of anyprolonged Earth orbital environment mustbe determined prior to manned operationof the Lab. The FCS must be capable ofstabilizing the Lab for docking with theCSM.

2.2.2.3 Earth Orbit Rendezous, Saturn IB Launch, Mission Profile (200 N.Mi, 200deg. Incl.)

3- Demonstrate unmanned LabFCS ability to operate afterorbital storage and effecta docking with the CSM.

MISSION PHASE NOMINAL PHASE TIME

Saturn IB'Ascent to Orbit(Lab)a) S-IB firingb) S-IVB firingc) LES boilerplate

CSM jettison

Earth Orbit Coast, S-IVB'Stabilizationa) Subsystems status check

(Hr: Min: Sec)

00:12:00

03:00:00


Demonstrate integratedmission control. (Note:Applicable during allphases through CSM-Labdocking.)

Demonstrate operation ofmodified LEM subsystemsafter being subjectedto boost environment.

S-IVB-Lab Separationa) SLA petal deployb) IMP initiated sepc) IMP shuts down S/S

Earth Orbit Storage(Lab unstabilized)a) IMP activates S/S

for status checks

00:20:00

Variable (Fromseveral hr up to30 days.)

None

Demonstrate ability of LabECS, EPS, and Inst & Commequipment to perform duringunmanned orbital storage.

(XVI)2-7

MISSION PHASE

5- Saturn IB Ascent to Orbit(GSM & Crew)a) Same as la & b, aboveb) LES jettison

6. Earth Orbit GSM C/0

7. CSM-SIVB Separation

8. GSM Rendezvous with Lab

9. GSM Active Dockinga) IMP stabilizes Labb) CSM docks

10. Lab C/0a) Crew transferb) Lab C/0

11. Earth Orbit LabExperimentOperations:a) Biomedical &

behavioral studiesb) EVA studiesc) Radiation monitoring

NOMINAL PHASE TIME(Hr: Min: Sec)

00:12:00

12. Lab Shutdown and Equipmentand Data Transfer to CSM

13. Preparation for CM Entrya) Lab jettisonb) SPS deorbit maneuverc) SM jettison

14. Entry

15. Parachute Descent

16. Post Landing ThroughS/C Retrieval

00:30:00

00:05:00

Variable

00:15:00

01:30:00

336:00:00(lit- days)

01:30:00

00:10:00

00:11:00

00:07:00

Variable


None

None

None

None

Demonstrate unmanned LabFCS ability to operate afterorbital storage, and effectdocking with a manned CSM.

None

Evaluate the:• EVA crew operations.• Lab data management

system.• Capability of modified

LEM ECS & crew provisionsto support life.

• Performance of modifiedLEM structure and insula-tion over a 14-day mission.

• Pointing capability of theLab FCS.

(XVI)2-8

2.2.3 Earth Polar Orbit, Saturn V "Launch


The launch configuration for all Saturn V launched AES missions will be the sameas the basic Apollo missions, and consists of S-IC, S-II and S-IVB booster stages,GSM and Lab, SLA, and LES. For the polar mission trajectory, the booster will belaunched from KSC in a south easterly direction (l44 deg azimuth), with the space-craft inserted directly into a 200-n.mi, 90-deg inclination orbit, using a yawsteering maneuver prior to orbital insertion. Experiment operations will becarried on by a three-man crew for 14 days. The CM will return the crew for awater landing and recovery in the Pacific recovery area.

2.2-3.2 Spacecraft Test Objectives

A polar mission profile adds the following spacecraft test objectives to the basicobjectives listed "in Paragraph 2.2.1.

OBJECTIVE

1. Demonstrate Lab structuralintegrity after Saturn Vyaw steering maneuver toeffect a polar orbit.

RATIONALE

1. Based on the current launch schedule, thisis the first Lab launched using the SaturnV booster. Although the Saturn V environ-ment will be known, the added effect ofthe booster yaw steering maneuver must bedemonstrated.

2. Demonstrate compatibilityof Lab with the MSFN forpolar orbit mission opera-tions .

3- Evaluate radiation levelsthroughout missionduration.

2. The present MSFW stations are designedfor low-inclination Earth orbit trajec-tory coverage. The capability of thepresent and new MSFN stations to monitorand track .the spacecraft must be deter-mined.

This is the first manned Earth polarmission. Cosmic ray intensity is greatestnear the poles and the resulting effectson equipment and crew must be studied.

2.2.3.3 Earth Polar Orbit, Saturn V Launch, Mission Profile (200-W.Mi, 90-DegInclination)

MISSION PHASE

Saturn V Ascent to PolarOrbita) S-IC firingb) S-II firing (yaw

steering)c) LES jettisond) S-IVB firing (yaw

steering)

NOMINAL PHASE TIME(Hr: Min: Sec:)

00:16:30


Demonstrate Lab structuralintegrity after Saturn Vyaw steering maneuvering toeffect a polar orbit.

(XVI)2-9

MISSION PHASE NOMINAL PHASE TIME(Hr: Min: Sec:)

2.

3.

Orbit Coast, S-IVB Stabilized 03:00:00a) GSM C/0b) GSM transpositionc) CSM-LEM Lab dockingd) Lab S-IVB separation

Lab C/0a) Crew enters Labb) Activation & C/0 of

Lab subsystems

Earth Polar Orbital LabExperiment Operations:a) Biomedical & behav-

ioral studiesb) Lunar survey equipment

c/o (Earth mapping)

02:00:00

336:00:00(Ik days)

5. Lab shutdown & equipmentand data transfer to GSM.

6. Preparation for CM re-entrya) Lab jettisonb) Service propulsion

deorbit maneuverc) SM jettison

7. Entry

8. Parachute descent

9. Post landing throughS/C retrieval

00:30:00

00:10:00

00:11:00

00:07:00

Variable


None

Demonstrate performance ofmodified LEM subsystemsafter being subjected toboost environment.

Demonstrate the:• Pointing capability of

Lab.FCS.• Performance of the

modified LEM structure &insulation over a 14-dayEarth orbital mission.

• Capability of the modifiedLEM ECS and-crew pro^visions to support lifeover a lU-day Earthorbital mission.

• Compatibility of Labwith MSFN for polarorbital operations.

Evaluate the:• EVA crew operations.• Radiation levels through-

out mission duration.• Lab data management sys-

tem.

None

(XVI)2-10

2.2.4 Earth Synchronous Orbit, Saturn V Launch


The Saturn V launch vehicle is used for manned synchronous earth orbital missions.The spacecraft including the S-IVB stage is first inserted into a 100-n. mi, 28.5-deg inclination Earth orbit. The S-IVB is restarted at the fourth descendingnodal crossing for a partial plane change and a simultaneous transfer to synchro-nous orbit altitude of 19,350 n.mi. The SPS and S-IVB circularize and completethe plane changes necessary to maintain a stationary lU-day orbit. Ground stationcoverage is continuous for this mission and all experiments may be ground monitoredand/or recorded in real time. Experiments and orbital operations will be carriedout by a three-man crew. Radiation monitoring will be of primary importance onthis type mission because of prolonged operations within the Earth radiation belts.


The Earth synchronous orbit mission profile adds the following spacecraft objec-tives to the basic objectives listed in Paragraph 2.2.1.

1.

OBJECTIVES

Demonstrate Lab thermalcontrol during synchro-nous orbit mission profile.

2. Demonstrate compatibilityof Lab with the MSFW forsynchronous orbit missionoperations.

RATIONALE

1. The synchronous mission profile, in con-junction with a 14-day mission durationhas not been demonstrated during thebasic Apollo missions. The thermalenvironment differs from the low Earthorbit in that the Earth re-radiation isreduced, and the spacecraft remains indirect sunlight for longer periods oftime. In addition, several experimentsplace attitude constraints on the Labwhich preclude maintaining a thermalbalance by rotating the spacecraft as isdone in the Apollo missions. The effi-ciency of the Lab thermal control systemshould therefore be demonstrated on thistype mission prior to a lunar .orbitingmission which has similar attitude con-straints imposed by the experiments.

2. Since only two synchronous missions areplanned, the first manned flight mustconfirm ground communications capability.The MSFE must have the capability of

. providing continuous coverage foroperations required to obtain synchronousaltitude, and for operations while inorbit. Ground station coverage mustalso be available for the return portionof the mission.

(XVI)2-11

OBJECTIVE RATIONALE

3. Evaluate radiation levelsthroughout missionduration.

3. This is the first manned synchronousorbit mission. The effect of pro-longed exposure to both cosmic andouter Van Allen belt radiation oncrew and electronic components mustbe determined. The effectiveness ofcrew and equipment shielding must bedemonstrated.

2.2A.3 Earth Synchronous Orbit, Saturn V Launch, Mission Profile (19350 N.Mi,

0° Incl.)

MISSION PHASE

1. Saturn V Ascent to 100-N.MiParking Orbit 28.5-DegInclina) Launch azimuth 90 degb) S-IC firingc) S-II firingd) LES jettison.e) S-IVB firingf) Parking orbit insertion

2.-. Earth Parking .Orbit (3 Orbits)a) -GSM systems checkb) Preparation for transfer

orbit insertion

3. Transfer Orbit Insertiona) S-IVB restartb) Orbit transfer & partial

plane change

4. Transfer Orbit Coast to19,350 n.mi

5-. S-IIB Restart to doPartial Plane Changeand Orbit Change atApogee

6. Elliptical Orbit Coasta) CSM S-IVB separationb) CSM transposition &

docking with Labc) Lab S-IVB separationd) Orbit coast to apogee


00:12:00

05:00:00

00:06:00

05:15:00

00:02:00

13: 1:00

OPERATIONAL TESTOBJECTIVE

None

None

None

None

None

None

(XVI)2-12

MISSION PHASE

7. GSM Firing to CompletePlane Change and Circulari-zation

8. Lab C/0a) Crew transfer to Labb) Activation and C/0 of

Lab S/S

9. Earth Synch Orbit LabExperiment Operationsoperationsa) Biomedical & behavioral

studiesb) Astronomical studies and

observationsc) Small maneuverable

satellite studies


00:01:30

01:30:00

336:00:00(Ik days)

10. Lab shutdown & equipmentand data transfer to CSM.

11. Preparation for CM re-entrya) Lab jettisonb) Service propulsion


12. Entry


lh. Post Landing through S/Cretrieval

01:30:00

05:30:00

00:11:00

00:07:00

Variable


'None

Demonstrate opera-tion of modified •LEiM S/S after beingsubjected to. boostenvironments.

Demonstrate the:• Lab thermal

control.• Compatibility

of Lab withMSFW for synchorbit mission.

• Capability ofthe modifiedLEM ECS andcrew provisionsfor l4-day mis-sion.

• Pointing capa-bility of LabFCS.

Evaluate the:• Radiation

levels• EVA crew opera-

tions .• Lab data man-

agement systems.

Demonstrate per-formance of modi-fied LEM structureand insulation.

(XVI)2-13

2.2.5 Lunar Orbit, Saturn V Launch

2.2.5.1 .Mission Description

The launch, parking orbit insertion, trans-lunar insertion, trans-lunar coast, andlunar orbit insertion will be identical to the Apollo lunar mission. The trans-earth insertion, trans-earth coast, re-entry, and recovery will also be the same.The vehicle will orbit the moon in an 80-n.mi. altitude orbit with a 10-deginclination. Earth ground station coverage is continous during line-of-sightoperation. This mission is designed to obtain detailed lunar mapping data in thevisual, IE, and radar frequency ranges. Visual observations will be made on thelight-side of the Moon, while radar and IE surveys will be made of both sides. Thismission will utilize the lunar survey equipment developed on an earlier AES earthorbital mission.


The following spacecraft test objectives are applicable to lunar orbital missions.The basic Phase I Lab test objectives of Paragraph 2.2.1 are accomplished duringearlier Phase I flights. .

OBJECTIVE

1. Evaluate radiation and micro-meteorite levels in low-inclination lunar orbitalenvironment.

Demonstrate CSM/LEM Lab Imission compatibility forlunar orbital thermalvacuum environment overmission duration.

RATIONALE

1. This is the first Lab long-duration lunarorbital mission. The radiation levelsthat can be expected on a lunar landingmission and their resulting effects onspacecraft electronics and crew shouldbe determined for a long-duration lunarmission.

2. This is the first lunar orbital missionwith the modified LEM configuration.

. Although the configuration will have beenflown in long-duration Earth orbitalmissions, the effects of thermal vacuumin the lunar orbital environment must be •confirmed. The spacecraft will be re-quired to point at the lunar surface forlong periods of time to accomplish mappingand survey tasks. This will limit atti-tude changes for thermal control. Theability of the modified LEM ECS to main-tain crew comfort and supply adequatevehicle cooling during these phases mustbe shown. The ability of the modifiedLEM structure and insulation to surviveboost environments and to maintain integrityin a lunar thermal vacuum environment mustbe demonstrated.

(XVl)2-lU

2.2.5.3 Lunar Orbit, Saturn V Launch, Mission Profile

NOMINAL PHASE TIMEMISSION PHASE

1. Saturn V Ascent to ParkingOrbit (100 n.mi.)a) S-IC firingb) S-II firingc) LES jettisond) S-IVB firinge) Parking orbit insertion

2.- Earth Parking Orbita) GSM C/0b) Preparation for trans -

lunar insertion

3. Translunar Insertiona) S-IVB restart

k. Translunar Coasta) CSM-S-IVB separationb) Transposition & dockingc) . GSM/Lab-S-IVB separationd) 1st mid-course correctione) 2nd mid-course correctionf) 3rd mid-course correction

5. Lunar Orbit Insertiona) SPS firing

6. Lab C/0 in Lunar Orbita) Crew transfer to Labb) .Activate & C/0 of

Lab systems

7. Lunar Orbit LabExperiment. Operationsa) Biomedical & behavioral

studiesb) Lunar photographic studiesc) Lunar oriented experiments

(Hr: Min: Sec)

00:12:00

02: 9:00

00:05:00

61:10:00

00:05:00

01:30:00

168:00:0.0(7 days)


None

None

None

None

None

Demonstrate opera-tion of modifiedLEM systems afterboost environments.

Evaluate radiationand micrometeoritelevels in low-inclination lunarorbit.Demonstrate CSM-LEM Lab I missioncompatability forlunar missionthermal vacuum en-vironment .

(XVI)2-15

MISSION PHASE

8. Lab shutdown and equip-ment and data transferto CSM.

9- Transearth Injectiona) Lab jettisonb) SPS firing


01:30:00

00:05:00

10. Transearth Coast 89:00:00a) 1st mid-course correctionb) 2nd mid-course correctionc) 3rd mid-course correction

11. SPS Deorbit 00:05:00a) SPS firingb) SM jettison

12. Entry ' 00:11:00

13. Parachute descent 00:07:00

'14. Post-landing through VariableS/C retrieval


None

None

None

(XVI)2-16

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TABLE 2.1-4

AES FLIGHT MISSION TEST PREREQUISITES

For Earth-Orbital AES Missions

The Apollo spacecraft shall have completed all Earth orbital flight development required for demonstrating lunar landing mission capability before engaging in AESEarth orbital missions. Specifically,:

• Block 11 CSM capability, including transposition, docking, Ik-day missioncapability, CSM life support, entry, Earth landing, rescue, and abort ofLEM crew shall have been demonstrated.

• LEM capability, including full duration, manned operation, propulsion andflight control operations, and subsystem performance in Earth orbit shallhave been demonstrated.

• MSFN/MCC operational'support for all phases of Earth-orbit missions andrecovery shall have been demonstrated.

• A representative amount of flight data using both Saturn IB and Saturn Vlaunch vehicles shall have" been acquired, analyzed, and compared with sub-system specification and performance data to assure Apollo capability forAES.

For Lunar AES Missions

The Apollo spacecraft shall have demonstrated lunar mission capability before en-gaging in AES lunar missions. Specifically,:

• Apollo expendable utilization and lunar mission time-line verification forboth subsystem performance and flight operational tasks .shall have beendemonstrated.

• Lunar surface characteristics, lunar landing and lunar take-off shall havebeen confirmed by a vehicle exhibiting LEM characteristics; notably, mass,inertia, center of gravity, gear tread, and crew support.

• MSFW/MCC operational support of lunar landing missions and recovery shallhave been demonstrated.

(XVl)2-23

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2.3 PHASE II LAB MISSIONS

2.3.1 Earth Orbit, Low-Inclination, Saturn V

2.3-1-1 Mission Description

The Phase II Lab Saturn V launched Earth orbital mission description is identicalto that of the Phase I Saturn V launched Earth orbital mission (Paragraph 2.2.1.1)However, the mission duration is 5 days. The Saturn V is capable of placing alarger payload in low Earth orbit than the Saturn IB. The prime purpose of thismission is to conduct extensive experiments throughout the increased missionduration consistent with the Phase II Lab capability. The prime test goal of thismission is to confirm the capability of the modified LEM subsystems to performsatisfactorily over a long-duration Earth orbital mission.


The following spacecraft test objectives are applicable to all Phase II Labmissions.

1.

2.

OBJECTIVE

Demonstrate performance ofthe modified.LEM structureand insulation over a 5-day Earth orbital Labmission.

Demonstrate operation ofmodified LEM subsystemsafter being subjected toboost environments.

Demonstrate the capabilityof the redesigned LEM ECSto maintain thermal con-trol over ^5-day earthorbital mission.

RATIONALE

1. The basic Apollo LEM is designed for a k-day active lifetime. The modifications tothe structure caused by equipment changesand experiment payload plus the modifiedinsulation must be evaluated in a long-

. duration thermal vacuum environment. Long-duration exposure to thermal vacuum environ-ment could cause structural or materialdegradation. Micrometeorite penetrationcould cause compromise of thermal shieldingand/or load bearing structures.

2. The Phase II Lab flights represent the firsttime that the modified equipment and instal-lation will be subjected to the boost environ-ment, and to the frequency variations andamplifications caused by the attachment ofthe new vehicle primary and secondarystructure. The flight results will verifythat the modified subsystem equipment willoperate after exposure to the launch environ-ment for the successful completion of thisand all subsequent Phase II Lab missions.

3. The earlier Phase I Labs will have used amodified LEM ECS for shorter durationmissions. To accomplish a 5-day mission,the ECS will be further modified and a per-formance demonstration on the first longduration mission is necessary. In addition,information on the functioning of the ECS

(XVl)2-25

OBJECTIVE

3. (Cont'd)

4.

8.

Demonstrate the capabilityof the two-gas ARS to sup-port the crew over a 45-day Earth orbital mission.

Demonstrate the performanceof Lab airlock and airlockcompatibility with vehicledesign and the two-gas ARS.

Demonstrate cryogenicstorage, distribution,and pressure regulationthroughout the 45-daymission duration.

Evaluate the operationalcharacteristics of theLab RCS in the spaceenvironment over the45 day mission.

Demonstrate satisfactoryoperation of the fuel cellEPS configuration in thespace environment over the45-day mission.

RATIONALE

equipment cooling loop (radiator, HpO boiler)is necessary to differentiate between elec-tronic subsystem and ECS cooling loop mal-functions in the event of thermal controlproblems.

4. The two-gas ARS (?0$> 02> 30% N2) will not havebeen flight tested previously. The abilityof the ARS to maintain proper mixture ratiosand proper pressure must be shown. The onlyway to confirm the redesigned ARS is toinvestigate it after exposure to launchaccelerations and the space environment overthe 45-day mission period.

5. The Phase II Lab flights will be the firsttime that an airlock is used for egress andingress. The ability to egress and ingresswithout depressurizing and pressurizing theentire cabin vill mean a savings in con-sumable usage. The compatibility of theairlock with the Lab design, so that pressureand gas leakage rates are minimized, must bedemonstrated. The ease with which the air-lock can be used must also be demonstrated.

6. The ability of the cryogenic tank heaters tomaintain the 0 and H? in the supercriticalregion must be shown. The zero-g effects onheater cycling and consequent pressureregulation must be shown. The fuel cellreactant supply system operation for a long-duration space mission must be shown.

7. The ability to start, restart, and cycle theRCS quads in the vacuum and the cold of spacewill be evaluated. Potential problems, suchas freezing or sluggish propellants, can beevaluated over a 45-day space mission.The effect on engine life of long-durationusage and outgassing of the thermal coatingof the thrust chambers under vacuum, hightemperature (due to usage) shock, andvibration will be evaluated.

8. Although fuel cells will have been flownpreviously, the overall Lab power generationsection including the fuel cells will nothave been exposed to a 45-day space oper-ational environment. The susceptibility offailure of the fuel cells and the cryogenicsupply system under prolonged zero-g effects

(XVI)2-26

OBJECTIVE

8. (Cont'd)

9- Demonstrate operation ofthe modified instrumenta-tion and data managementsystems.

RATIONALE

- will not be known because these types offailures are design dependent. Fuel cellcooling and electrolyte stratification arealso affected by the zero-g environment.This could cause the loss of all fuel cellsbecause of a common fault. The effectivenessof the fuel cell radiators to maintaintemperature must be evaluated.

The effectiveness of the modified PCM andthe modified GSM recorders in handling largeamounts of spacecraft and experimental datamust be demonstrated. The ability to dumplarge amounts of data in the limited groundstation coverage time available must bedemonstrated (to9..6-kb/s dump rate).

2.3.1.3 Earth Orbit, Saturn V Launch, Mission Profile (200 N. Mi)

NOMINAL PHASE TIME(Hr; Min; Sec)

00:12:00

MISSION PHASE

1. .Saturn V Ascent to Orbita) S-IC firingb) S-II firingc) LES jettisond) S-IVB firing

2. S-IVB Stabilization 03:00:00a) CSM C/0

3. Transposition & Docking 00:30:00a) CSM-S-IVB separationb) Transpositionc) CSM-LEM Lab dockingd) Lab extraction

from S-IVB

4. Lab C/0 01:30:00a) Crew transfer to Labb) Lab C/0 '

5. Earth Orbit Lab Experiment 1080:00:00Operations (45 days)a) Earth oriented applica-

tions


None

None

None

Demonstrate the:e Operation of modified LEM

subsystems after being sub-jected to boost environment.

e Capability of the redesignedECS to maintain thermal con-trol over a V?-day Earthorbital mission.

(XVI)2-27


8.

9-

10.

Lab shutdown and 01:30:00equipment and datatransfer to GSM

Preparation for CM 00:10:00entrya) Lab jettison

. b). SPS deorbit maneuverc) SM jettison

Entry 00:11:00

Parachute descent 00:07:00

Post landing through VariableS/C retrieval


Satisfactory operation offuel cells in space environ-ment over lj-5-day mission.Cryogenic storage, dis-tribution and pressureregulation throughout l»-5-day mission.Performance of modified LEMstructure and insulationover ^5-day Earth orbitalmission.Capability of the two-gasARS to support the crewover a U5-day Earth orbitalmission.Performance of Lab airlockand airlock compatibilitywith vehicle design and thetwo-gas ARS.Operation of the modifiedinstrumentation and datamanagement systems.Evaluate operational charac-teristics of the Lab ECS inspace environment over ^5-day mission.

(XVI)2-28

2.3.2 Earth Orbit Rendezvous, Saturn IB Launch


The Phase II Lab Saturn IB launched missions are the same as those of Phase I, i.e.,they are part of dual launch rendezvous and/or resupply missions. The Phase IILab mission duration is U5 days.


All Spacecraft Test Objectives for Phase I Earth orbit rendezvous, Saturn IBlaunch (Paragraph 2.2.2.2.), as well as the Phase II Lab Spacecraft Test Objectives(Paragraph 2.3.1.2) are applicable.

2.3.2.3 Earth Orbit Rendezous, Saturn IB Launch, Mission Profile (200 N.Mi, 28.5Deg Incl.)

MISSION PHASE

1. Saturn IB Ascent toOrbit (Lab)a) S-IB firingb) S-IVB firingc) LES-Boilerplate

GSM Jettison

2. Earth Orbit Coast S-IVBStabilizationa) Subsystems status

check

3. S-IVB - Lab Separationa) SLA petal deployb) IMP initiates sepc) LMP shuts down S/S

4. Earth Orbit Storage (LabUnstabilized)a) LMP activates S/S

for status C/0

5. Saturn IB Ascent toOrbit (GSM & Crew)a) Same as ~L, above

6. Earth Orbit GSM C/0

7. GSM - S-IVB Separation

8. GSM Rendezvous with Lab


00:12:00

03:00:00

00:20:00

Variable upto 30 days

00:12:00

00:30:00

00:05:00

Variable

OPERATIONAL TEST OBJECTIVES

Demonstrate integrated missioncontrol. (Note: Applicableduring all phases throughGSM-Lab docking.)

Demonstrate operation of modi-fied LEM subsystems after beingsubjected to boost environment.

None

Demonstrate ability of Lab ECS,EPS, Instr. & Comm. equipmentto perform during unmannedorbital storage.

(XVI)2-29

MISSION PHASE NOMINAL PHASE TIME(Hr: Min; Sec)


9.

10.

11.

GSM Active Dockinga) IMP stabilizes Labb) GSM docks

Lab C/0a) Crew transferb) Lab C/0

Earth Orbit Lab Experi-ment Operationsa) , Biomedical &

behavioral studiesb) EVA studiesc) Radiation monitoring

00:15:00

01:30:00

1080:00:00(45 days)

12. Lab Shutdown andEquipment and DataTransfer to GSM

13. Preparation for CM .Entrya) Lab jettisonb) SPS deorbit maneuverc) SM jettison

±k. Entry"


16. Post Landing throughS/C Retrieval.

01:30:00

00:10:00

00:11:00

00:07:00

Variable

Demonstrate unmanned Lab FCSability to operate afterorbital storage and effectdocking -with a manned CSM.

None

Demonstrate the:• Satisfactory operation of

the fuel cell EPS configu-ration over a Vj-day mission.

• Capability of the redesignedECS to maintain thermal con-trol over a 45-day mission.

• Performance of modified LEMstructure & insulation over45-day mission.

• Cryogenic storage, dis-tribution, and pressure reg-ulation throughout a 45-daymission.

• Capability of the two-gasARS. - .

• Performance of the airlock. '• Operation of the modified

instrumentation and datamanagement systems.

Evaluate operational character-istics of the Lab RCS over a45-day mission.

None

(XVI)2-30

2.3-3 Earth Polar Orbit, Saturn V Launch


The Phase II polar Earth orbit mission descriptions are the same as those of PhaseI Labs missions. Hov/ever} the mission duration is 45 days.


All Spacecraft Test Objectives for Phase I polar Earth orbit mission (Paragraph2.2.3.2), as well as the Phase II Lab Spacecraft Test Objectives (Paragraph2.3.1.2) are applicable.

2.3.3.3 Earth Polar Orbit, Saturn V Launch, Mission Profile (200 N.Mi, 90-DegInclination)

MISSION PHASE

1. Saturn V Ascent toPolar Orbita) S-IC firingb) S-II firing (yaw

steering)c) LES jettisond) S-IVB firing (yaw

steering)

2. Orbit Coast - S-IVBStabilizeda) GSM C/0

3. GSM Transpositionand Dockinga) GSM - SIVB

separationb) GSM transpositionc) CSM-LEM Lab/Dockingd) LAB-S-IVB

separation

14-. Lab C/0a) Crew enters Labb) Activation & C/0

of Lab subsystems

5. Earth Polar Orbital LabOperationsa) Biomedical &

behavioral studies


00:16:30

03:00:00

00:30:00

02:00:00

1080:00:00(45 days)


Demonstrate Lab structuralintegrity after Saturn V yawsteering maneuvering to effecta polar orbit.

Demonstrate performance of mod-ified LEM subsystems after beingsubjected to boost environment.

Demonstrate the:• Satisfactory operation of the

fuel cell EPS configuration -over a 45-day mission.

(XVI)2-31

MISSION PHASE

b) Weather studiesc) Geophysical

observations


6. Lab Shutdown andEquipment and DataTransfer to GSM

7. Preparation for CM Re-entrya) Lab jettisonb) Service propulsion


8. Entry


10. Post Landing throughS/C Retrieval

01:30:00

00:10:00

00:11:00

00:07:00

Variable


• Cryogenic storage, dis-tribution, and pressure reg-ulation throughout a U5-daymission.

• Performance of the modifiedLEM structure and insulationover a lj-5-day mission.

• Capability of the redesignedECS to maintain thermal con-trol over a ^5-day mission.

• Compatibility of Lab withMSFN for polar orbit oper-ations.

• Capability of the two-gasARS.

• Performance of the Labairlock.

• Operation of the modifiedinstrumentation and datamanagement systems.

Evaluate the:• Radiations levels, through-

out mission duration.• Operational characteristics

of the Lab RCS over a k1?-day mission.

None

(XVI)2-32

2.3.4 Earth Synchronous Orbit, Saturn, V Launch

2.3-4.1 Mission Description

The Phase II synchronous Earth orbit mission descriptions are the same as those ofPhase I Lab missions. The mission duration is 45 days.

2.3. .2 Spacecraft Test Objectives

All Spacecraft Test Objectives for Phase I synchronous Earth orbit mission (Para-graph 2.2.4.1), as well as the Phase II Spacecraft Test Objectives (Paragraph2.3.1.2) are applicable.

2.3-4.3 Earth Synchronous Orbit, Saturn V Launch, Mission Profile (19,350 N.Mi,0-Deg Inclination) .

MISSION PHASE

1. Saturn V Ascent to100-N.Mi ParkingOrbit, 28.5-deg inclin.a) Launch azimuth 90 degb) S-IC firingc) S-II firingd) LES jettisone) S-IVB firingf) Parking orbit

insertion

2. Earth Parking Orbit(3-1/4 orbits)a) CSM systems checkb) Preparation for

transfer orbitinsertion


00:12:00

05:00:00


None

None

3. Transfer Orbit Insertion 00:06:00a) S-IVB restartb) Orbit transfer &

partial plane change

4. Transfer Orbit Coast 05:15:00to 19,350 N.Mi

5. S-IVB Restart to do 00:02:00Partial Plane Change& Orbit Change atApogee.

None

None

None

(XVl)2-33

MISSION PHASE

6. Elliptical Orbit Coasta) CSM - S-IVB

separationb) CSM transposition

& docking with Labc) Lab - S-IVB

separationd) orbit coast

7. CSM firing to completeplane change and cir-cularization

8. Lab C/0a) Crew Transfer to

Labb) Activation and C/0

of Lab S/S

9. Earth Synch. Orbit LabExperiment Operationsa) Biomedical &

behavioral studiesb) Astronomical studies

and observations


13: 1:00


None

00:01:30

01:30:00

1080:00:00days)

Hone

• Demonstrate operation ofmodified LEM subsystemsafter being subjected toboost environments

Demonstrate the: •• Lab thermal control

throughout mission.• Capability of Lab with

MSFN for. synch, orbitmission. • • • • .

• Performance of the modifiedLEM structure and -insulationover a 5-day mission.

• Satisfactory operation of •• the fuel cell EPS con-figuration- over a ^5-day •mission.

• Cryogenic storage, dis-tribution and pressure reg-ulation through a 5-daymission. •_

• Capability of .the redesignedECS to maintain thermal con-trol over a k^-day mission.

• Capability of the two-gasARS.

• Performance of the Lab air-lock.

• Operation of the modifiedinstrumentation and datamanagement systems.

Evaluate the:• Radiation levels.• Operational characteristics

of the Lab RCS over a lj-5-day mission.

(XVI)2-314

MISSION PHASE

10. LEM Lab Shutdown& Equipment and DataTransfer to GSM

11. Preparation for CM Re-entrya) Lab jettisonb) Service propulsion


12. Entry

13. Parachute descent

14. Post-Landing throughSpacecraft Retrieval

NOMINAL PHASE TIME(Hr; Min: Sec)

01:30:00

05:30:00


None

00:11:00

00:07:00

Variable

2.3-5 High-Inclination Lunar Orbit, Saturn V Launch

2.3.5-1 Mission Description

These missions are similar to the Phase I lunar orbit mission; however,'the missionduration is 35 days, and are lunar polar orbital. The orbit is inclined 90 deg atan altitude of 80 n.mi. During lunar orbit insertion extra A V is applied by theservice module propulsion system (SPS), to give polar inclination.


The following spacecraft test objectives are applicable to lunar orbit missions.The Phase II Test Objectives (Paragraph 2.3.1.2) are considered accomplished duringearlier Earth orbital flights.

OBJECTIVE

1. Evaluate radiation andmicrometeorite levelsin high-inclination lunarorbital environment.

RATIONALE

1. Previous lunar orbital missions have evaluatedradiation levels and micrometeorite activityin low-inclination orbits. These phenomenamust be investigated during a long-duration,35-day, Phase II mission to show the effective-ness of the modified micrometeorite shielding,as well as the radiation effects on electronicsand crew.

(XVl)2-35

3.

OBJECTIVE -

Demonstrate CSM-LEM LabII mission compatibilityfor lunar orbital thermalvacuum environment overmission duration 35 days;

Demonstrate Lab IIstructural and systemintegrity after lunarpolar orbit insertionSPS firing.

Demonstrate Lab II ex-pendable utilizationduring extended lunarmission time line.

RATIONALE

2. This is the first long-duration.lunar orbitalmission in a high-inclination orbit.- Thethermal vacuum effects on vehicle structureand subsystem electronic components must bedemonstrated. The ability of the Phase IILab ECS to maintain crew comfort and performequipment thermal control must be demonstrated.

3. The effect of long-duration SPS firing onmodified subsystemsafter- translunar coast ina thermal vacuum environment must be demon-strated. The negative-g effects on modifiedstructure and subsystems must be demonstrated.

k. Expendable usage for long-duration lunarmissions will not have been demonstratedpreviously. Although long-duration, earthorbital mission, consumable usage for PhaseII Labs will be known; the different environ-'mental and mission conditions require ex-pendable usage to be demonstrated in lunarorbit. •

2.3.5.3 Lunar Orbit, Saturn V Launch, Mission Profile

MISSION PHASE

1. Saturn V Ascent to•Parking Orbit (100 n.mi)a) S-IC firingb) S-II firingc) LES jettisond) S-IVB firinge) Parking orbit

insertion

2. Earth Parking Orbita) CSM C/0b) Preparation for

translunar insertion


h. Translunar Coasta) CSM-SIVB separationb) Transposition &

dockingc) CSM/Lab-SIVB

separation

NOMINAL PHASE TIME(Hr; Min"; Sec)

00:12:00


Rone

02: 9:00 None

00:05:00

61:10:00

None

None

(XVl)2-36



5.

6.

7.

8.

9.

10.

11.

d) 1st .mid-coursecorrection .

e) 2nd mid-coursecorrection

f). 3rd mid-coursecorrection •

Lunar Orbit Insertiona) SPS firing for orbit

insertion.b) SPS firing plane change.

Lab C/0 in Lunar Orbita) :Crew transfer.to Labb) Activate & C/0 of Lab

systems

Lunar Orbit Lab•Experiment Operationsa) Biomedical &

behavioral studiesb) Lunar photographic

studiesc) . Lunar oriented ex-

periments

00:05:00

01:30:00

672:00:00(28 .days)

LEM Lab Shutdown andEquipment and DataTransfer to CSM

Transearth Injectiona) Lab jettisonb) SPS firing

Transearth Coasta) 1st mid-course

correctionb) 2nd mid-course

correctionc) 3rd mid-course

correction

SPS Deorbita) SPS firingb) SM jettison

01:30:00

00:05:00

89:00:00

00:05:00

Demonstrate Lab Structuraland system integrity afterlunar polar, orbit insertionSPS firing. •

Demonstrate operation of modi-fied subsystems after boostenvironments.

Evaluate radiation and micro-meteorite levels in high-in-clination lunar orbit.-Demonstrate the:o CSM-LEM Lab II mission

compatibility for lunar .mission thermal vacuum en-vironment, 35 days,

o Lab II expendable usageduring extended lunarmission stay time.

None

(XVI)2-37

2.k SHELTER MISSIONS

2.U.I Mission Description

The Shelter launch, earth parking orbit, translunar insertion, transluar coast,and lunar orbit insertion mission phases are the same as an Apollo lunar mission.The. Shelter is dormant before separation from the GSM in lunar orbit. After lunarorbit checkout, the Shelter separates unmanned and performs a pre-programmed descentto a preselected site. Two" methods of lunar descent are being considered. Method Ais a direct descent method from an initial orbit of 20 n.mi. Method B is an Apollolunar mission Hohrnann'transfer descent from 80 n.mi. The non-essential Shelter sub-systems are shut down after touch-down and the CSM returns to earth. The Shelterhas the capability of a 90-day stay time on the lunar surface. Ground stationcoverage is continuous for line-of-sight portions of the mission. The prime purposeof the Shelter is to provide a base for lunar'surface exploration and experiments.All three planned Shelter missions are basically identical. Mission variations willdepend on the Shelter payload.

2.U.2 Spacecraft Test Objectives

The following spacecraft test objectives are applicable to the Shelter missions.The Shelter spacecraft test objectives will complement the spacecraft tests andinvestigation performed during the Phase I and early Phase II Lab missions.

1.

OBJECTIVE

Demonstrate Earth-Shelter communi-cation 'capability.during lunar orbitand descent to.the lunar surface.

2. Demonstrate unmanned pre-programmedinertial landing system.

Demonstrate landing stability andstructural integrity of theShelter/payload combination.

RATIONALE

1. The ability of the Coram -Subsystemto transmit data to the Earth duringlunar orbit operations and line-of- -sight descent must be demonstrated.The subsystem capability to transmitvia VHF to the CSM during non7line-of -sight operations and to switch todirect communications with the Earthmust be demonstrated.

2. This is the first time that an un-manned lunar landing will have beenattempted. The ability of the ShelterGW&C to maintain spacecraft controlduring descent engine thrusting andto automatically follow a pre-programmed descent profile to a pre-selected sight must be demonstrated.The ability of the 8-jet RCS tomaintain vehicle control must bedemonstrated.

3. Although Apollo lunar landings willhave been performed, the effects ofthe modified structure and subsystemsplus a large payload on landingstability must be shown. The effects

(XVI)2-38

OBJECTIVE

Demonstrate the performance of theShelter thermal control -designduring quiescent stay time.

Demonstrate fuel cell section opera-tional capability' from cold storagethrough warmup and earth commandactivation.

6. Demonstrate operation of modifiedsubsystems after being subjected toboost environment.

RATIONALE

of landing loads during an unmannedlanding on vehicle structural integ-rity must be demonstrated.

The capability of the Shelter ECS tomaintain .thermal control during, theunmanned lunar storage period mustbe demonstrated. Phase I and IILab's ability to maintain thermalcontrol during extended and unmannedEarth orbital storage periods willhave been demonstrated. However,this capability in a lunar environ-ment has not been shown before.Addition of the. RTG affects, thethermal control. The effectivenessof the modified thermal insulationand of additional equipment heatersmust be demonstrated. The heatrejection ability of the thermalradiators must be demonstrated.

The ability to maintain fuel cellreactant tank pressures withinacceptable limits during lunar •storage must be demonstrated. Thecapability of the fuel cells to warm'up and then be activated by groundcommand must be shown. Fuel cellstartup after thermal vacuum storagewill not have been demonstratedpreviously.

The LEM subsystems will have beenflown on both the Saturn IB andSaturn V boosters. Modifications tothe structure and supports requirethat loads and.vibrations impartedto -the spacecraft by the launch vehi-cle and by the Shelter to the variousequipments and equipment areas beconfirmed. Major additions to th>3 EOSand EPS (RTG, fuel cells and associ-ated tankage) plus deletion of theascent propulsion subsystem and 50$of the RCS cluster hardware requirethat the boost environment effectson the overall Shelter be confirmed.

(XVI)2-39

Demonstrate the performance ofthe Shelter airlock during multiplelunar surface egress and ingresscycles.

The ability of the astronaut toeffectively use the airlock wearinga hardsuit and handling equipmentmust be demonstrated. The abilityto maintain a shirtsleeve cabinatmosphere during egress and ingresscycles must be shown. The pressureand atmosphere leakage rates of thecabin to the airlock must be deter-mined.

2.4.3 Shelter, Saturn V Launch, Mission Profile

MISSION PHASE

1. Saturn V Ascent to ParkingOrbit (100 n.mi)a) S-IC firingb) S-II firingc) LES jettisond) S-IVB firinge) Parking orbit insertion

2. Earth Parking Orbita) CSM C/0b) Preparation for trans-

lunar insertion


k, Translunar Coasta) CSM-SIVB separationb) Transposition & dockingc) GSM/Shelter - S-IVB

separationd) 1st mid-course

correctione) 2nd mid-course

correctionf) 3rd mid-course

correction

5. Lunar Orbit Insertiona) SPS firing (Note: 20-n.

mi circular for directdescent, 8o-n. mi cir-cular for Apollodescent.)

NOMINAL PHASE TIME(Hr: Min: Sec:)

00:12:00


None

02: 9:00 None

00:05:00

61:10:00

None

None

00:05:00 None

(XVI)2-UO

MISSION PHASE

6. Shelter C/0 and LunarOrbit.. ' - -a) Crew transfer to

.; Shelterb) Activate-So C/0

Shelter systems

NOMINAL PHASE TIME(Hr; Min;' Sec:) ,

03: 3:00(80-n.rai orbit)

02: 9:00(20-n.mi orbit)


Demonstrate the:• Earth-Shelter communi-

cations capabilityduring lunar orbit.

• Operation of .modified. subsystems after beingsubjected to boostenvironment.

Shelter-CSM Separationa) __ Preparation for

separationb) Remove inhibit from

descent programmerc) Crew return to CSMd) Shelter separation

Lunar Descent & LandingMethod A - Direct Descentfrom 20 n.mia) Initiate powered descentb) Braking phasec) Final approach phased) Landing phase

Method B - Hohmann transferfrom'80 n.mia) Transfer orbit insertionb) Initiate powered descent

at 50,000 ftc) Braking phased) Final approach phasee) Landing phase

Lunar Storage Perioda) Post landing S/S C/0b) S/S shutdownc) Periodic status checks

00:20:00 None

00:15:00

01:10:00

(up to 90 days)

10. Lunar Shelter Manned . 336:00:00Experiment Operations (lU days)a) Topographical studiesb) Geological studiesc) Photography & radiometryd) Gravity and magnetic studiese) Extra vehicular activities

11. Shelter Shutdown & Equip- 01:30:00ment Transfer to Taxi

Demonstrate the:• Unmanned preprogrammed

inertial landing system• Earth-Shelter communi-

cation capability duringdescent to the moon.

Demonstrate landing sta-bility and structuralintegrity of the Shelter/payload combination.

Demonstrate performance ofthe Shelter thermal controlduring quiescent stay time.

Demonstrate the:• Fuel cell section opera-

tional capability fromcold storage throughwarm-up and Earth commandactivation.

• Performance of the Shelterairlock during multiplelunar surface egress andingress cycles.

None

(XVI) 2-1

MISSION PHASE NOMINAL PHASE TIME OPERATIONAL TEST OBJECTIVES(Hr; Min: Sec)

12. Entry 00:11:00

13. Parachute Descent 00:07:00

14. Post-Landing through VariableS/C Retrieval.

(XVI)2-U2

2.5 TAXI MISSIONS - '

2.5.1 Mission Description

The Taxi launch, earth parking orbit, translunar injection, translunar coast, lunarorbit insertion, lunar orbit checkout, transfer orbit insertion, and lunar descentand landing are the same as that of an Apollo lunar landing mission. The mannedTaxi lands at the selected location and the crew transfers to the Shelter. At thecompletion of the lunar liftoff, GSM docking and transearth injection sequences fol-low the basic Apollo mission. Taxi launch occurs prior to GSM passage overhead.Directly over launch site, at an altitude of 50,000 ft, burnout occurs followed bya 3/ij- orbit rendezvous coast. The Taxi rendezvous and docks with the GSM and datatransfer. Approximately 1-3A orbits after Taxi/GSM docking, transearth injectionis initiated. The GSM initiates a multi-impulse or single-impulse transearth traj-ectory maneuver. The return to Earth is the same as an Apollo lunar mission. TheCSM-Taxi mission duration is 23 days, of which ik are on the lunar surface. Groundstation coverage is continuous for all line-of-sight operations. The prime purposeof the Taxi is to provide a means for crew arrival and departure from the lunar sur-face. All three planned Taxi mission are the same as the basic Taxi mission des-cribed.

2.5.2 Spacecraft Test Objectives

The .following spacecraft test objectives are applicable to all Taxi missions. TheTaxi spacecraft test objectives will implement the spacecraft tests and investiga-tions performed on the Phase I and early Phase II Lab missions.

1.

OBJECTIVE

Demonstrate the performance ofthe Taxi thermal control design fora l4-day lunar stay mission.

RATIONALE

1.

2. Demonstrate Taxi landing with theaid of a transponder

Demonstrate the capability of theTaxi ECS and crew provisions tosupport crew in accordance with theTaxi mission profile.

The additions and modifications'tosubsystem equipment, plus the addi-tion of thermal insulation, windowcovers, and top docking tunnel coverrequire that the ECS's ability tomaintain thermal control of cabinand equipment be demonstrated. Theeffectiveness of such heat transferequipment as the RTG "heat pipe"must be shown during the lunar stay

i quiescent period.

2. This is the first time a lunar land-ing using a beacon (on Shelter) forhoming purposes has been planned.The ability of the rendezvous radarto acquire the beacon and use it forlunar descent and landing in vicinityof the Shelter must .be demonstrated.

3. The ability of the Taxi ECS to main-tain thermal control of cabin andequipment and to maintain cabin pres-sure and 02 supply must be shown. Theability of the ECS to perform repeatedpressurization/depressurization/pres-surization cycles must be shown.

(XVI)2-li3

2.5-3 Taxi, Saturn V Launch, Mission Profile

NOMINAL PHASE TIME(Hr; Min: Sec:)

00:12:00

MISSION PHASE

1. Saturn V Ascent toParking orbit (100 n.mia) S-IC firingb) S-II firingc) LES jettisond) S-IVB firinge) Parking orbit

insertion

2. Earth Parking Orbit 2: 9:00a) CSM C/0b) Preparation for trans-

lunar insertion

3. Translunar Insertion 00:05:00a) S-IVB restart

. -Translunar Coast 6l:10:00a) CSM-S-IVB separationb) Transposition & dockingc) CSM/Shelter-S-IVB

separationd) 1st mid-course

correctione) 2nd mid-course

correctionf) 3**d mid-course

correction

5. Lunar Orbit Insertion 00:05:00(80 n.mi)a) SPS firing

6. Taxi C/0 and Lunar Orbit 03: 3:00a) Crew transfer to Taxib) Activate and C/0 Taxi

systems

7. Taxi-CSM Separation 00:20:00a) Preparation for

separationb) Taxi active separation

8. Lunar Descent and Landing 01:10:00a) Transfer orbit

insertionb) Initiate powered descent

at pericynthianc) Braking phase


None

None

None

None

None

None

None

Demonstrate the:• Taxi landing with aid

of a transponder.• Capability of the Taxi

ECS and crew provisionsto support crew in

MISSION PHASE

d) Final approach phasee) Landing phasef) Post landing C/0

9. Taxi Storage Perioda) Shutdown of Taxib) Crew transfer to

Shelterc) Shelter operations

10. Taxi Ascenta) Systems activation,

and Prelaunch C/0b) Ascent

11. Taxi-CSM Rendezvousa) Taxi active rendezvousb) Taxi active docking

•12. Taxi Shutdown & Equipmentand Data Transfer to CSM

13. Transearth Injectiona) Taxi jettisonb) SPS firing

lU. Transearth Coasta) 1st mid-course

correctionb) 2nd mid-course

correctionc) 3rd mid-course

correction

15. SPS Deorbita) SPS firingb) SM jettison

16. Entry


18. Post-Landing through .S/C Retrieval -

NOMINAL PHASE TIME'(Hr: Min: Sec:)

336:00:00(Ih days)

01: 0:00

00: 5:00

01:30:00

00:05:00

89:00:00

00:05:00

00:11:00

00:07:00

Variable

OPERATIONAL TEST'OBJECTIVE

accordance with the Taximission profile.

Demonstrate the:• Performance of Taxi

thermal control designduring 1^-day mission.

• Capability of the taxiECS and crew'provisionsto support crew inaccordance with theTaxi mission profile.

• Capability of the' TaxiECS and crew provisionsto support crew in.accordance with theTaxi mission profile.

(XVI)2-1»5

2.6 TRACKING REQUIREMENTS

This section presents the operational ground tracking coverage and communicationsrequirements for AES missions. A brief summary of the MSFN capabilities is given.Subsequently, the results of an analysis of the ground station coverage times,applicable to the Earth orbital and lunar missions, are summarized. Finally, theLab, Shelter, and Taxi operational communications requirements are described andcompared with MSFN capabilities.

2.6.1 Ground Rules

The tracking requirements are based on the following:

• The projected Apollo Manned Spaceflight Network (MSFN) of 11 near-space andthree deep-space 'stations will be utilized for AES missions.

2.6.2 Assumptions

The tracking requirements are based on the following assumptions:

• For comparison between spacecraft communication requirements and MSFNcapabilities, it is assumed that given four distinct S-Band frequencies,two may be processed and recorded in real-time and the remaining twosignals are recorded.

• S-Band ground antenna line-of-sight is limited to 5 deg above the horizon.• For lunar missions coverage, time begins upon lunar insertion.• Parking-orbit insertion is taken as mission initiation.• Thrusting maneuvers, involving Hohmann transfer from parking-orbit to a

higher operational altitude, are not considered. . Parking-orbit altitudeis assumed for the duration of transfer resulting in less coverage time.

• All orbits are considered circular.• Tracking ships provide continuous coverage for earth-orbital-missions up to

point of insertion.

2.6.3 Summary of MSFN Capabilities •

The projected Apollo Unified S-Band (USB) system will consist of 11 near-spacelocations (30-ft antenna), deployed for Earth-orbital and lunar-injection phasecommunications, along with three deep-space sites (85-ft antenna) providing coverageat translunar and lunar distances. For ranges greater than 10,000 n. mi (a nominalboundary differentiating near-space and deep-space active tracking), deep-spacestations perform as active transmitting and receiving stations, while near-spacesites passively track. This technique enables a multi-station "fix" on the space-craft's position. Table'2.6-1 designates locations, operational date and near-spaceor deep-space mode of operation. Also tabulated is site "dual" or "single" capa-bility as described below.

2.6.3.1 Dual Stations

Eight stations, including the deep-space locations, will have "dual" capabilitiespermitting concurrent tracking of two space vehicles within the antenna main beamwidth (85-ft antenna: 0.35 ± 0.05 deg, 30-ft antenna: 1.05 ± 0.25 deg). Duringthe dual track mode at lunar distances (ground antenna pointing is performed onone spacecraft), a minimum beamwidth of 0.30 deg limits the separation between

(XVI)2-U6

spacecraft, in a plane perpendicular to an earth station, to approximately 500 n.mi (Fig. 2.6-1). Cognizance of such an existing constraint, although not expectedto be exceeded during any. phase of the lunar missions, should, nonetheless, betaken. . -'

The deep- space stations will have facilities necessary for real-time processing oftwo S-Band signals. In addition, two other frequencies may be received; however,this information will be recorded. According to References U-2 and U-3, threeS-Band frequencies are the maximum number of signals which would be received attranslunar and lunar distances; two telemetry signals could be processed in real-timeand the third recorded. Further inquiry into ground station facilities impliesthat four S-Band frequencies could be received, hence it has been assumed thatthis additional signal could only be recorded. Coverage afforded by the deep- spacestations at translunar and lunar distances is continuous in that at least one sitewill always be in "view" of the spacecraft (2 -hour coverage, clearly presupposesno lunar eclipsing of the spacecraft). Furthermore, there are certain periods whentwo locations can maintain communications with the spacecraft.

2.6.3.2 Single Station

The remaining five stations will possess "single" -capabilities for communicationsand tracking support of one spacecraft. These stations will be able to receiveand process two distinct S-Band frequencies.

2.6.U Ground Tracking Analysis for Earth-Orbital Missions

The Phase I Lab missions discussed in Paragraph 2.2 may be categorized into fivegroups, according to 'earth' or lunar orbital altitude and inclination. This para-graph presents a summary of ground station tracking coverage time available for theih Unified S-Band locations. Eleven near-space MSFN stations are primarilyconsidered for low- inclination earth-orbital missions. In addition to theselocations, the effects of introducing the Fairbanks, Alaska site for polar orbitshas been examined. The three deep- space MSFN locations provide coverage for lunarorbital missions, with the near-space sites as back-up.

2. 6. U.I Low- inclination Earth-Orbit

The mission profile is 200-n. mi altitude and 28.5-d.eg inclination. The averagecoverage per orbit is 19-73

A graphical representation of the first four orbits is shown in Fig. 2.6-2.Figure 2.6-3 gives a ground coverage analysis for the initial 36 orbits. Individualstation coverage time is shown along with the total time available for tracking perorbit (station overlap included). Orbits 11 through 13 are indicative of an orbitalgrouping where minimum coverage is afforded. Also arising, is the possibility of noground tracking, typified by orbit 27. It is possible that by planning non- criticalmission events, during the orbits where no coverage is afforded this situation, -maybe tolerable. Thirty-six orbits are shown in Fig. 2.6-3, however, approximaterepetition was found to occur after Vf orbits. This periodicity is typical of all200 n.mi altitude missions.

(XVI)2-U7

2.6.4.2 Polar Earth-Orbit

The mission profile is 200-n.mi altitude and 90-deg inclination. The averagecoverage per orbit is 8.43 min for USB sites, and 11. 14 min when FBKS is added.

The ground track of the initial 4 1/4 orbits of a polar mission is shown inFig. 2.6-4. The apparent discontinuity arises from the characteristics of theMercator projection in that the polar ground track exhibits the same disjointednessas the Meridians. Table 2.6-2 indicates coverage available for the first thirty-six orbits. Total coverage per orbit is given first for the projected USB stationsalone and then for the case where the Fairbanks, Alaska location is included.

In conclusion, implementing USB at the Fairbanks location would be highly desirableas indicated by the increased coverage time. This analysis indicates that utiliz-ation of this site extends the average orbital coverage by 32$. An additional im-portant result is the elimination of orbits without ground contact.

2.6.4.3 Polar Retrograde-Orbit

The mission profile is 200-n.mi altitude and -83 deg (retrograde- orbit) inclination.The average coverage per orbit is 8.4-5 min for USB sites, and 11.30 min when FBKSis added.

Shown in Fig. 2.6-5 is the ground track of the initial 5 1/4 orbits, and Table 2.6-2indicates coverage available for the first 36 orbits. -Total coverage per orbit isgiven for USB stations alone and for ttie case when Fairbanks, Alaska location isincluded.

In conclusion, implementing USB at the Fairbanks location would result in similarbenefits as in the previously cited Polar Earth-Orbit mission.

2.6.404 Synchronous Earth-Orbit

The mission profile is 19,350-n.mi altitude and 0-deg inclination,, The coverage iscontinuous. Continuous coverage is possible in synchronous orbit since the orbitalperiod coincides with Earth rotation, hence the spacecraft appears virtuallystationary.

Figure 2.6-6 illustrates the entire ground track sequence until synchronous orbitalinsertion. Figure 2.6-7 shows the ground station coverage from .parking orbit in-sertion, through the Hohmann transfer to synchronous" altitude.

In conclusion, the ground station visibility chart, Fig. 2.6-7, shows that thespacecraft is continously tracked from a point 1 hr after transfer orbit insertionto synchronous altitude. At the synchronous point (0° lat., 168° W long.) thespacecraft is visible to seven ground stations, simultaneously.

2.6.5 Ground Tracking Analysis for Lunar Missions

2.6.5.1 Low-Inclination Lunar Orbit (80-n.mi Altitude)

Figure 2.6-8 graphically represents the coverage times afforded by the three active(transmitting and receiving) ground stations for lunar orbits „ The solid barsindicate the spacecraft is ground- tracked, whereas the cross-hatching signifies

(XVI)2-U8

lunar eclipsing of the vehicle. The upper time scale designates hours elapsed sub-sequent to lunar insertion, while the lower scale identifies the day. Nearly13-2/3 days were taken before repition due to the lunar cycle occurred. 13-2/3days constitutes one-half of the complete lunar cycle. The next half-cycle repre-sents the mirror image. And every 13-2/3-day interval will have a similar pattern.However, due to the Earth rotation, the ground station tracking sequence may bedifferent. This periodicity is typical of all lunar-orbital and lunar-stay missions.

Table 2.6-3 summarizes the ground station coverage times for the following threeperiods: (l) Total 13-2/3-Day Interval, (2) Continuous Coverage, (3) LunarOccultation. Period 1 is the entire ground track time shown in Fig. 2.6-8. Thisperiod is further subdivided into three tracking modes. Column one indicates thetotal time available for tracking by one station. The second column gives thetotal coverage time afforded by two stations simultaneously. Summation of the firstand second yields the total effective coverage and is shown in third column. Period2 is the time interval providing continuous spacecraft coverage. The second periodis subdivided into three tracking modes as follows: Average Coverage - One Station;Average Coverage - Two Stations; Total Average Coverage. The third period accountsfor all the time intervals characterized by lunar eclipsing of the spacecraft. Sub-division of this period coincides with the continuous period cited above.

Table 2.6-3 indicates no continuous coverage periods for the low-inclination typemissions. The spacecraft is in ground station "view" approximately 1 hr and 17 minper orbit, while the average occultation period lasts h6 min.

On an average, the lunar orbiting vehicle is ground-tracked 65^ of the time. Thespacecraft will spend approximately 2/3 of the total coverage time in "view" of oneground station. The remaining time will be available for simultaneous ground-tracking by two stations.

2.6.5.2 Lunar Polar-Orbit

Figure 2.6-9 charts the ground coverage time for the point of lunar insertion untilthe lunar cycle repeats 13-2/3 days later. A summary of. the results appears inTable 2.6-3. (The description of the figures and the tabulation is the same as.Paragraph 2.6.5.1.)

In Fig. 2.6-9, hours 37-125 illustrate a continuous coverage period exhibiting nolunar interposition. Hours 0-37 and. 125-328 represent time intervals marked bylunar occultations. Such a pattern, characterized by alternate intervals ofcontinuous and intermittent coverage periods, is unique for the high-inclination,lunar-orbital missions. Figure 2.6-10 depicts the effects of lunar position onearth coverage. Position 1 illustrates partial lunar eclipsing while position 2shows no lunar interference.

The results shown in Table 2.6-3 indicate that coverage (by either one or two deep-space stations) is available j6% of. the time. Furthermore, spacecraft "viewing"time is approximately doubled in the single tracking station mode as compared tothe double station mode. The second column of Table 2.6-3 (Continuous CoveragePeriod) suggests a contradiction; however this is due to an averaging over a short-time interval when the-most favorable double station coverage exists.

(XVI) 2-1

2.6.5.3 Shelter or Taxi Lunar Stay

A satisfactory approximation of the ground tracking capability for the Shelter/Taxilunar stay coverage times may be taken from the low-inclination, lunar-orbital.,mission tracking analysis. Therefore, Fig. 2.6-8 is applicable if the cross-hatchedlunar occultation periods are assumed to be solid bars, signifying ground coverage.(For a description of the ground coverage figures and the tabulated summary referto the discussion on the Low-Inclination Lunar Orbit, Paragraph 2.6.5.1.)

As shown in Table 2.6-3, the tracking time spent in line-of-sight of one groundstation exceeds that of the two-station mode by two-fold. Ground coverage will becontinuous with the possible exceptionoof two small gaps (Fig. 2.6-9, hours 26^ and288). The maximum time interval between coverage periods is nominally 1^ min. Thegaps are neglected at the present time since this duration is small and their oc-currence infrequent.

2.6.6 Lunar Mission Communications Operations

2.6.6.1 Phase I - Earth-Orbit

Station coverage for launch, Earth-orbital, and translunar injection phases aresimilar for all Apollo lunar missions. Information concerning ground-tracking timefor parking orbit may be found in Paragraph 2.6. .1, noting that the 100-n.miinitial earth-orbit, typical for lunar missions, results in less coverage time thanindicated for 200-n.mi earth-orbital flights. Ground Station tracking will be pro-vided by the MSFW near-space stations for this phase.

Communications between Shelter, Taxi, or Lab (depending upon mission) and Earthduring the earth-orbital phase are not required; however, GSM data acquisition andmonitoring will be necessary during this period. A single phase modulated rf signal,transmitted by the GSM, provides the data link. CSM FM transmission is possible ona second rf carrier; therefore, all .USB stations may receive and process both sig-nals in real-time, simultaneously.

2.6.6.2 Phase II - Translunar

During the translunar phase, continuous telemetry and tracking coverage is affordedby the three deep-space stations for all AES missions. Spacecraft communicationrequirements are identical to Phase I (Paragraph 2.6.6.1). Ground capability in-creases to the extent that all active locations possess "dual" capability.

2.6.6.3 Phase III - Lunar

2.6.6.3-1 Lab (Lunar Orbit). While 2k-laour coverage is available throughout the trans-lunar flight, lunar orbiting introduces occultation periods, where the spacecraft,shielded by lunar interposition, is not within line-of-sight. This periodic lossof communication will necessitate planning of real-time scientific experiments totake place directly after and prior to lunar interference. Experiments capable ofbeing recorded may be performed at any time and "dumped" once earth-spacecraft ac-quisition is established.

The CSM-Lab combination will not separate during the lunar-orbital phase of themission, thus excluding the requirement of tracking two spacecraft for the lunar-orbital Lab missions. Communication requirements, as far as the number of carrier

(XVI)2-50

frequencies needed, are similar to Phases I and II herein. Digital informationtelemetered to Earth, however, increases in conjunction with the scientific experi-ment "being performed and has a maximum rate of 09.6 kbps.

2.6.6.3.2 Shelter - Direct Descent. Presently, there are two methods of Shelterdescent under discussion. One suggests a direct lunar landing while the othermethod coincides with the basic Apollo descent. This paragraph describes thedirect landing, while the Hohmann transfer method is summarized in a followingparagraph which discusses the Taxi (or Shelter) coast and descent.

Since abort capabilities are not necessary for the unmanned Shelter, a direct lunardescent method is possible. The mission description may be summarized as: Lunarinsertion begins while the Shelter-CSM combination is eclipsed by the moon, followedby a circular orbital coast at an altitude of 20 n.mi, nearly 0-deg inclination.Approximately 1-1/2 orbits later, Shelter separation and direct descent is initiated.Separation and subsequent powered landing from the coasting orbit could possibly beperformed while in line-of-sight of the CSM and Earth, for landing sites in thewestern longitudes. The two proposed lunar landing locations are Alphonsus explora-tion site (13° 13'S, 1° 20'W) and Hyginus Crater (7° Vp'N, 6° l8'E). Figure 2.6-11illustrates the preceding mission sequences.

Visual monitoring of the unmanned Shelter during descent is performed by the astro-nauts in the CSM. Full Shelter status data will be transmitted (51-2 kbps) directlyto Earth until successful landing and storage is verified. Continuous Earth-tracking coverage of both the CSM and Shelter vehicles is provided by the deep-space stations.

During this period there exists the possibility of three different S-Band frequen-cies being transmitted simultaneously. Two frequencies originating from the CSM,one carrying communications and another TV, while a third gives Shelter trackingand status data. Present deep-space station capabilities, when in view of onestation, allow for concurrent real-time processing of two signals while the thirdsignal, if necessary, may be recorded. If complete real-time'processing of allthree frequencies is required, then this phase must be carried out in view of twodeep-space stations.

2.6.6.3.3 Shelter - Lunar Stay. During the Shelter storage mode, prior to Taxilaunch, status data will be transmitted periodically to Earth in real-time, low- orhigh-bit rate upon uplink command. Forty-eight hrs preceding the Taxi launch, fullShelter status data will be telemetered and shall continue until Taxi arrival.Earth stations shall be able to provide 24-hour, real-time coverage.

2.6.6.3.4 Taxi - Lunar Insertion, Coast and Descent (Alternate Shelter Method).In lunar vicinity, the mission description coincides with the basic Apollo flightand is summarized as follows: Lunar orbital insertion will begin out of line-of-sight of Earth, followed by an 80-n.mi, low-inclination coast of one orbit. Sepa-ration of the Taxi-CSM occurs on the second orbit. The Hohmann descent transferto 50,000 ft is initiated approximately 20 min later. 'One-quarter of an orbitlater, the CSM-Taxi spacecraft emerges from the backside of the moon into line-of-sight of earth stations. Figure 2.6-11 shows lunar insertion along with theHohmann descent transfer.

The Taxi-CSM separation and the initial phase of the Hohmann descent transfer bothoccur during lunar occultation resulting in loss of communication with Earth

(XVI)2-51

stations. The Taxi status, during this period, will be transmitted to the CSM(VHF, 1.6 kbps) and recorded. Upon acquisition of the earth stations, the CSM willtransmit the recorded data to earth. Data transmission at 51.2 kbps when Earth-Taxi acquisition and lock-on is established. At this time, Taxi-CSM communicationscease and a direct Earth-Taxi link is maintained. Deep-space station capabilitiespermit full tracking of both CSM and Taxi when in line-of-sight of the Earthantenna.

Utilization of all S-Band carriers implies that a total of four frequencies will benecessary during Taxi descent phase. Frequency allocation proceeds as follows:CSM, requiring two carriers for communications and TV; Shelter, telemetering fullstatus data; Taxi communications. Considering singular deep-space station coverage,two S-Band frequencies may be processed in real-time. The remaining two signalscan be recorded. For complete real-time processing support, the descent phase mustoccur during a period where two deep-space station coverage is available. If thetwo-station coverage method is not feasible, then station augmentation is required.

2.6.6.3.5 Taxi - Lunar Landing. The Taxi-powered landing from 50,000 ft to touch-down occurs while in line-of-sight of the CSM, Earth, and the Shelter. A Shelter-Taxi transponder link provides a guide for the Taxi landing.

Earth-spacecraft communication requirements are identical to those of the LunarInsertion, Coast and Descent (Paragraph 2.6.6.3. ).

2.6.6.3-6 Taxi - Lunar Stay. In addition to the aforementioned S-Band frequenciesrequired, the mobile payloads, i.e., the LSSM and MFS, operating independently ofthe Shelter-Taxi, will utilize one S-Band frequency each, thus increasing the totalnumber of signals received by Earth to six. Although continuous coverage will beafforded, the possibility of six simultaneous S-Band frequencies must be considered.

Assuming deep-space station support is not augmented, then it would be necessary toimplement such procedures as: time-sharing of signals; scheduling communicationswhen two-station coverage is available. The following suggests a possible Earth-Moon communications method. During the lunar stay, the Shelter will accommodateboth astronauts while the Taxi, in the storage mode, provides status transmissionupon earth command via an onboard timer sequencer. Periodic monitoring of the Taxican be performed while the CSM, in lunar orbit, is eclipsed by the moon. Upon LSSMor MFS deployment, astronaut A, located in the Shelter, will monitor astronaut B,conducting exploration. This precludes the possibility of simultaneous operation ofboth LSSM and MFS. Some communications possibilities for ground support are givenin Table 2.6-1*.

This table is divided into two major columns indicating whether the CSM is shieldedby the Moon or is in line-of-sight of the Earth. Only one operational mode isshown in the first major column, whereby real-time support is extended to the mannedvehicles (Shelter and MFS/LSSM). The CSM transmission frequencies (fg, f ) are notrepresented in the first major column (CSM Eclipsed by Moon) since no Earth-GSM icommunication link exists.

The second major column of Table 2.6-4 (CSM in LOS of Earth) is further subdividedinto two modes of communications. In both modes the LSSM or MFS is assumed to beoperational; however, only in Mode 1 is there real-time support. Real-time com-munication, with Earth is maintained between manned Shelter and Earth for both modes.Note that the Taxi transmission frequency (fi) is not shown in either mode.

(XVI)2-52

In the preceding discussion it was assumed that ideal communications existed betweenground stations and vehicles located on the lunar surface. The crater-marked lunarsurface suggests that such a terrain would reduce coverage afforded by groundstations. That this is the case for the L3SM/MFS and not for the Shelter/Taxi willbe described below.

Study of the two lunar landing sites (Alphonsus and Hyginus Crater) indicates thatthere are no immediate obstructions to limit communications for the Shelter/Taxi.The lunar terrain does introduce an approximate 6-deg mask angle above the horizonin certain directions; however, this will not reduce Earth-tracking coverage. Aproblem arises in the case of MFS or IfiSM communication. In this case the mobilevehicles may at times be in such a position where higher terrain may directly blockEarth from view, resulting in a loss of communication. (See Vol. V, Sec 5 5 3)

TABLE 2.6-1

(XVI) 2- 53M

Unified S-Band Station Capabilities

Station

Carnarvon

Bermuda

Hawaii

Cape Kennedy

Texas (Corpus Christi)

Guaymas

Guam

Ascension

Antigua

Grand Canary Is .

Grand Bahama Is .

Golds tone

Canberra

Madrid

NearSpaceSite

X

X

X

X

X

X

X

X

X

X

X

DeepSpaceSite

X

X

X

DualorSingle

D

S

D

D

S

S

D

D

S

S

S

D

D

D

ProposedOperationalDate

Dec '66

Jan '67

Feb '67

Mar '67

May '67

Apr '67

Jan '67

Jun '67

Oct '67

Sep '67

Jul '67

Apr '67

Jul '67

Oct '67

Ref: SID 6 -1866

Rev

123456789

101112IS141516IT181920212223242526272829303132333536

A. 83-<3eg Inclination

Site Coverage Time, min*

TEX/7. 49,,GYM/5 . 48FBKS/6.96,FBKS/7.59,FBKS/5-42,ASC/4.09,ANT/6. 79 ,CRO/7.51,TEX/7.13,GYM/3.16,HAW/ 5. 24,HAW/ 5. 45Asc/i.63,ASC/6.34,CRO/2.50,CRO/6.78,GYM/7.43,FBKS/3.89FBKS/7.54,FBKS/7.15CYI/7.42,ASC/7.42,CRO/4.54,CRO/6.20,GYM/7.56,FBKS/7.73HAW/7.30,CYI/3.38GAM/1.72,GAM/6.48,CRO/6.69,CRO/6.69,GYM/7.29FBKS/7.73,CYI/5.53,

GYM/ 5. 79

HAW/4.65HAW/ 5. 88ASC/5 . 56

GAM/7.42BDA/7.08GBI/7-53,GYM/6.80,FBKS/7.56FBKS/7.04

CYI/7-56GAM/7.52BDA/7-38,CKEW/7.38,TEX/6.11

HAW/7.18

FBKS/3.95,GAM/5.62ANT/7.52,GBI/6.33,TEX/4.58,

FBKS/6.00

CYI/6.74ANT/4.09,BDA/7.42,BDA/7.42,

HAW/7.41FBKS/6.41

CYI/7 • 31

CKEN/7-57, BDA/2.79FBKS/5.28

AM/7. 47, GBI/1.00GBI/7.01, TEX/5.00

GAM/4.32

BDA/7.56, GBI/4.52, CKEN/3.5'CKEN/6.81, TEX/6.33, FBKS/l'.FBKS/6.37

ASC/6.67CKEN/5.62, GBI/6.47TEX/7.24, GBI/4.22

*Station sequence in chronological order

(XVl)2-55<

2.6-2 Site Coverage Times

>9

Total Cov'ge/rev,min

USBSites

7-495.1»84.655.8812.8711.5110.8015.667.633.165. 2k5.459-1913-8613-5014.337.43

7.18

11.7613.0215-9813.627-56

7-303-388.461 .8516.939-707.29

7.4l5-53

USB &FBKS

7.495.48ll.6l12.4718.2911.5110.8015.6612.9110.7212.315.459.1913-8613.5014.337.433.8914.727.1515.7113.0215.9814.7113-937-7313,303-388.4614.8516.939.707-295-9215.1411.94

B. 90-deg Inclination

Site Coverage Time, min*

TEX/6.94, GYM/6.97FBKS/5-97, GYM/1.78 •FBKS/7. 73, HAW/6.21 ,FBKS/6.01, HAW/3-71 1ASC/6.24, CYI/6.84 ;ASC/1.29, CYI/2.54, GAM/7.07ANT/6.46, BDA/6.3 :

CRO/7-55, GBI/7.55, ANT/1.77, CKEN/7-45, BDA/5.07TEX/7.43, GYM/6.06GYM/ 4. 87, FBKS/6.20HAW/4.17, FBKS/7. 72 !HAW/6.00, FBKS/ 5- 76ASC/1.90, CYI/7-21ASC/6.15, GAM/7.52CRO/5.69 BDA/7.54, GBI/5.47, CKEN/4.40, ANT/7.00CRO/5.04, CKEN/6.35, TEX/6.78, GBI/5.47GYM/7- 50, TEX/2. 91FBKS/7 • 10 ;FBKS/7.49, HAW/7.53FBKS/ 4. 02ABC/7. 51, CYT/7-53, GAM/6. 08GAM/2.93CRO/3-36, ANT/7.52, BDA/7-57, GBI/3-12CRO/6.52, GBI/6.68, CKEN/7.12, TEX/5- 72GYM/7.54, TEX/5.31, FBKS/2.28FBKS/7 • 21HAW/7.22, FBKS/7. 42FBKS/3.46, CYI/6.38 .ASC/6.97, GAM/2.47, CYI/4.26GAM/6 . 20 , BDA/ 5 . 44 ANT/6 . 31CRO/7.50, BDA/6.05, CKEN/7.24, GBI/7-46, ANT/2.64TEX/7.54, GYM/5-31FBKS/4.87, GYM/5.74FBKS/7. 64, HAW/3.16FBKS/6.75, HAW/6.37ASC/2.92, CYI/5-02

Total Cov ' ge/rev,min

USBSites

7.051.786.213-7113.0810.9010.3015.357.434.874.176.009.1113.6716.7012.007.50

7.53

21.122.938.8613.927-54

7.226.3813.5017.95.16.007.545.743.166.378.00

USB &FBKS

7.057-7513.949-7213.0810.9010.3015-357.4311.0711.8911.76 •9.1113.6716.7012.007-507.1015.024.0221.122.938.8613-929.827.2114.649.8413.5017-9516.007-5410.61.9.8013.128.00

. (XVl)2-57

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(XVI)2-59/60

A. Moon as Viewed fromLunar North Pole

I Distance Projected onto a Plane J.to Earth Station

B. Moon as Viewed from Earth

Fig. 2.6-1 Lunar Projected Planar Distance

j/tusnmcui-

ANT Antiqua

ASC Ascension

BDA Bermuda

CKEN Cape Kennedy

CRO Carnarvon

CYI Grand Canary Island

FBKS Fairbanks (Alaska)

GAM Guam

GBI Grand Bahama Island

GYM Guaymas

HAW Hawaii

TEX Texas (Corpus Christi)

Note: Station Coverage Shownfor 200 n mi

135 180

90

Longitude, deg

45 45 90

Fig. 2.6-2 Ground Track - 28.5 DegInclination Orbit

Station

TEXGYMHAWGAMCROASCCYIANTBDAGBICKEN

Time, hrrev


Time, hrrev

65


Time, hrrev

129

?<(,-

33

I97

~T1511

T14

10

^~(XVI)

Rev

1

2

3

4

\7Total x

Coverage/Rev

17.43

16.38

39.23

38.62

I5

5

6

7

8

17.90

13.52

10.79

8.33

I10 11

9

10

11

12

11.75

11.79

6.7

2.92

I16

I17

12Fig. 2.6-3 Vf9tfeet 1 of 3) Ground Station

$- Visibility, 28.5 deg InclinationOrbit; Alt: 200 n mi

Station


Time, hrrev

1813

I1!


Time, hrrev

2417

2S


Time, hr 30rev 21

14

I20

I2115

2618

2719

I32

22

I332"3

(xvi)2-65/66

Rev

13

14

15

16

TotalCoverage/Rev

7.10

21.23

24.44

29.85

I22

1623

28I29

20

17

18

19

20

27.89

33.31

36.22

36.76

21

22

23

24

19.45

11.96

7.19

4.22

3424

35

Fig. 2.6-3 (Sheet 2 of 3)

Station

TEXGYMHEGAMCROASCCYIANTBDAGBICKEN

Time, hr 36rev 25

137


Time, hr 42rev 29

143


Time, hr 48rev 33

49

I38

26

I44

30

I50

34

3927

I4531

I5135

(XVI)2-67/68

Rev

25

26

27

28

TotalCoverage/Rev

13.00

14.72

6.54

I40

28

I41

29

30

31

32

16.89

24.29

25.66

26.41

I46

3247

33

34

35

36

33.35

34.87

38.25

29.40

I52

I53

36

Fig. 2.6-3 (Sheet 3 of 3)

C

ANT Antiqua

ASC Ascension

BDA Bermuda

CKEN Cape Kennedy

CRO Carnarvon



GAM Guam


GYM Guaymas

HAW Hawaii



135 180

(XVI)2-6^/70

200 n .mi.Insertion

90Longitude, deg

o

60

45 90

Fig. 2.6-4 Ground Track - 90 DegInclination Orbit

ANT Antiqua

ASC Ascension

BDA Bermuda

CKEN Cape Kennedy

CRO Carnarvon



GAM Guam


GYM Guaymas

HAW Hawaii

T.EX Texas (Corpus Christi)


135 180

35 90

Longitude, deg

45 90

Fig. 2.6-5 Ground Track - 83 DegInclination Orbit

ANT Antiqua

ASC Ascension

BDA Bermuda

CKEN Cape Kennedy

CRO Carnarvon



GAM Guam


GYM Guaymas

HAW Hawaii


A Deep SpaceTracking Stations


135 180 135

90

Longitude, deg Fig. 2.6-6 Ground Track -Synchronous Orbit

Golds tone

Canberra

Madrid

TEX

GYM

HAW

GAM

CRO

ASC

CYI

ANT

BDA

GBI

CKEN

Time, hr 10 .First Apogee t19,350 n. mi ; — 'Jettison SIVB

11- f f - 15

:

3 ' BegiiTo

16 [ Transfer OrbitPerigee: 5,140n. mi

(XVI) 2-757/76'

Transfer 5ich Orbit —*

•4 J-18 20 23 Synchronous

Orbit24 Continuous

Coverage

Fig. 2.6\7 Ground Station Visibility -» \ Synchronous Orbit

MadridCanberraGoldstone

(HourTime]

I Day

Y///A


(HourTime]

I Day

48 56


I HourTime]

I Day


(HourTime]

iDay

96

144

104

152


(HourTimel

iDay

192 200


(HourTime!

I Day

240

10

248


(HourTime]

bay

288

12

296

LegendGround CoverageLunar Occultation (No Ground Coverage)

'///A

Y///A

Y//A

V//S,'/A

Y///A V/A4L

16 24i

'S///A

V////A

'///AV///AW//A V///A

64 72 81

mr/// V///A

'///A'/A

112 120

Y///A V///AV///A

160 168 1'

W//A

208 216

Y///M V//A'///A

V///A

256 264

'///A

11

V&7MV//A'///A

V/A

304 312

13

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96

'///s

136

'///A

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V///

184


144


192

24

72

20

232

280

End of Lunar Cycle

Repetition Begins

328


240

10


288

12


336

14

Fig. 2.6-8 Low-Inclination Lunar OrbitGround Station Visibility

MadridCanberraGolds tone

(Hour 0Time! *-

(Day 0


(Hour 48Time] ,_

Pay 2

56


(Hour 96Time | ,_

I Day 4

104


Time(Hour 144

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152


(Hour 192Time{ i—

I Day 8

200


(Hour 240Time{ i_

(Day 10


(Hour 288Time{ i

IDay 12

£2/2

248

Y///A

296

LegendGround CoverageLunar Occultation (No Ground Coverage)

16 24'

64 72

112 120

160 168

208 216

256 264

11

304 312

13

(XVI)2-T9/80

32

80

40

88

MadridCanberraGolds tone

48_ j

2


96

128 136

176 184


144

6


192

'///A

224

272

320

232

280

' End of Lunar Cycle-

Repetition Begins

328


240

10


288

12


336

14

Fig. 2.6-9 Lunar Polar Orbit GroundStation Visibility

• (XVI)2-81

Position 1

Portion of Polar OrbitShielded from Earth by Moon

Position 2Continuous Earth Coverageof Polar Orbit

Fig. 2.6-10 Effect of Lunar Position on Earth Coverage

(XVI)2-82

Out of GroundStation LOS

1. Portion of Trans-lunar Flight

2. Lunar Insertion3. Shelter Separation

& Direct DescentAfter 1.25 RevolutionCoast

Shelter

CSM

A. Shelter Coast and Landing, Direct Descent Method

1. Portion of Translunar Flight2. Lunar Insertion (80 n mi)3. Taxi-CSM Separation4. Initiation of Taxi Hohmann

Descent Transfer to 50, 000 ft5. Begin Powered Descent at

50,000f t6. Lunar Landing

TaxiCSM

*EarthB. Taxi Coast and Landing (Alternate Shelter Method)

1. Taxi Lift-Off2. Begin Coast for

CSM Docking3. Taxi/CSM Rendezvous

at 80 n mi4. Initiate Transearth

Injection5. Portion of

Transearth Flight

C. Taxi Ascent and Transearth Injection

Fig. 2.6-11. Shelter/Taxi Operations in Lunar Orbit

(XVI)2-83

2.7 OPERATIONAL DATA HANDLING

2.7.1 Assumptions

The operational data handling analysis is based on the following assumptions:

• The Apollo Ms si on Control Center (MCC) will be used as a base for AESmission control and data handling.

• All MSFN remote sites supporting AES flight operations will be remoted tothe MCC-Houston.

• Sufficient data will be routed to the MCC in real-time to conduct theflight control and experimental program.

• Ground control will retain the capability of acquiring and analyzingsufficient significant data so that specific contingencies or malfunctionscan be isolated and the crew advised that a potential need to modify orabort the flight may exist.

2.7.2 Basic Spacecraft Data Requirement

The mission control philosophy on which this analysis is based is that the AESmissions follow the basic Apollo LEM missions, except that more responsibility isplaced upon the crew for spacecraft subsystem monitoring. The objective of the MCCreal-time data monitoring is to provide assistance to the crew in achieving missionobjectives and recommend action in emergency situations. Table 2.7-1 presents apreliminary definition of operational subsystem and aeromedical parameters for real-time MCC display and compares these parameters to the onboard spacecraft displays.The parameter listing is based upon the recommended Phase I and II Labs, Shelter,and Taxi configurations (Vol III - VI). As subsystem design and aeromedical re-quirements become more detailed, and as the basic configurations reflect particularexperiment requirements, the data handling requirements will be modified and ex-panded accordingly.

2.7.3 Experiment Data Interface

Experiment data in the case of the Labs will not be transmitted to the ground inreal-time. Two on-board tape recorders will be used enabling continuous recordingcapability during lengthy experiments, and subsequent transmission to the groundfollowed by tape erasure. The magnitude of experiment data per mission, the mini-mal ground station coverage during near-earth and polar trajectory flights, and ex-periment packing per flight impose a requirement for quick tape verification by theacquiring ground station. This capability is requisite in deciding to erase thetape (required for continuous data acquisition during lengthy experiments and forproceeding to the next experiment), re-dumping when adequate ground station cover-age is available, or repeat the experiment (if deemed necessary because on-boardacquisition was not accomplished, etc,). Because of the need for expeditious con-firmation of tape verification, the read-after-write method is proposed. The useof read-after-write techniques would allow signal strength and quality to be deter-mined concurrently with the stripping out, buffering, and processing of significantparameters for display. These parameters would constitute a cross sample repre-'sentation of data tape content for tape verification. After satisfactory acquisition

(XVI) 2-81*

of the experiment data, the data should be expedited to MSC, Houston for disposi-tion.

2.J.4 Operational Data Flow .

The operational data flow should follow, as closely as possible, the currentlyplanned Apollo/tEM data flow. The operational telemetry data will be acquired bythe remoted MSFN site and formatted to the existing ground communications equip-ment. Near real-time processed data and summary messages will be selected fortransmission to MCC. MCC will retain and store the data that has been receivedand will be able to selectively recall, reduce, and/or display it on the displayconsoles or make available hard copy of particular data. When practicable or- whenrequested, MSFN will ship all of the operational data it has acquired (magnetictapes, strip chart records, etc.) to MSC for logging into the Control Metric DataFile. All reduceable data will then be processed, copied, and reduced per exist-ing or special data processing request. The processed data will aid mission analy-sis during and after the flight. The GAEC copies of the operational data will beexpedited to Grumman's facility. There it will be logged in, processed, reduced,and .analyzed for detailed subsystems performance evaluation. Operational data thatis required for correlation with experiment data will be recorded such that groupsof related operational and experimental data can be stripped-out with minimumcomplexity.

2.7.5 Instrumentation Subsystem

2.7.5-1 Onboard Instrumentation Subsystem Description

The Instrumentation Subsystem consists of operational equipment, caution and warn-ing equipment, and experiment support equipment. The major sections within thesubsystem are as follows:

• Operational Equipment Section; This section will detect, measure, andprovide to the on-board displays certain data for monitoring and evaluatingsubsystem performance. It will also provide data and "real-time" refer-ence for transmission to earth, to facilitate ground assessment of vehicleperformance and failure analysis. The operational instrumentation sectionwill consist of sensor equipment, signal conditioning equipment and pulsecode modulation and timing equipment (PCMTE).

• Caution and Warning Section; This section provides constant monitoring ofcertain critical subsystem areas and supplies displays and telemetry datawhen malfunctions occur.

• Experiment Support Equipment: This equipment consists of a LEM pulse codemodulation (PCM) unit and two modified Apollo tape recorders. Inputsignals to the equipment are pre-conditioned, and all experiments willprovide separate display outputs.

2,7.5.2 Measurements Lists

A preliminary measurements analysis has been made for the four basic types ofspacecraft, and the complete parameter tabulations are presented in the VehicleDesign Analysis Reports, Vol III through VI. These lists indicate how the measure-ments are made and monitored. Tables 2.7-2 through 2.7-5 summarize the total

(XVl)2-85

number and type of measurements available, per subsystem, that may be used tosatisfy both ground control and flight test objectives for the AES spacecraft.Experiment instrumentation measurement listing is beyond the scope of this report.

(~lAusn/na/L.

(XVI)2-86

Table 2.7-1

REAL-TIME STATUS DATA HANDLING REQUIREMENTS

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(XVI)2-87

Table 2.7-1 (Cont'd)

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1 Skin Temperature1 Radiator Outlet Temperature1 Output CurrentOutput .Voltage

FCA No.FCA No.FCA No.FCA No.FCA No.FCA No.FCA No.FCA No.FCA No.H20 No.FCA No.FCA No.FCA No.FCA No.FCA NoFCA NoFCA NoFCA NoFCA No

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Coolant Radiator Inlet TemperatureRTG TemperatureRTG VoltageRTG CurrentENVIRONMENTAL CONTROL SUBSYSCabin TemperatureCabin PressureAirlock PressureSuit Inlet TemperatureSuit Outlet PressureFCA H20 Tank QuantityHeat Pipe Condenser TemperatureHeat Pipe Evaporator TemperatureRadiator Outlet Glycol Temperature•Coolant Radiator TemperatureDescent H20 Tank No. 1 QuantityDescent H20 Tank No. 2 QuantityAscent H20 Tank No. 1 QuantityAscent H20 Tank No. 2 QuantityCoolant Pumps A PressureCoolant Accumlator Fluid Low LevelLiOH SEL/EE COgPLSS Supplement Tank PressureSuit Diverter VPI ClosedSelect H20 Separator RateH20 Separator No.. 2 SelectCC^ Cartridge in Secondary PositionC02 Partial PressureEmergency Og VPI OpenEmergency Og Valve Elect OpenPrimary Suit Compressor Failure

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Spare Suit Compressor FailureSelect Suit Compressor FailureCoolant Pump No. 1 FailureCoolant Pump Ho. 2 FailurePrimary Pump Discharge PressureRedundant Pump Discharge PressureSelected Pump Outlet PressureMain Water Boiler Coolant Cutlet TempRedundant Water Boiler Coolant Outlet TempSelected Water Boiler Coolant Outlet TempHgO Separator No. 1 RateHgO Separator No. 2 RateGUIDANCE, NAVIGATION & CONTROLComputer Digital DataIMU Standby/OffACA Out Of DetentDeadband SelectSCS Mode Select (Auto)SCS Mode Select (Att Hold)Pitch Trim FailureRoll Trim FailureLGC WarningInertial Reference WarningPGNS CautionAEA Test Mode FailureACS Power Supply FailureGuidance Select Switch - AGSACS Mode (Stand-By)AGS Mode (Warm-Up)Abort Stage CommandedASA TemperaturePIPA Temperature+ 32 VDC Abort Sensing Assembly+ 12 VDC Abort Sensing Assembly- 12 VDC Abort Sensing Assembly+ 6 VDC Abort Sensing Assembly+ It- VDC Abort Sensing Assembly- 2 VDC Abort Sensing Assembly+ 15 VDC Supply Voltage- 15 VDC Supply Voltage+ k VDC Supply Voltage- 4 VDC Supply Voltage+ 6 VDC Supply Voltage- 6 VDC Supply VoltagePitch RGA Signal (0.8kc) VoltageYaw RGA Signal (O.Skc) VoltageRoll RGA Signal (O.Skc) VoltageRGA Pick-Off Exct O.Skc VoltageRGA Spin Motor A Ph O.Skc VoltageRGA Spin Motor B Ph O.Skc VoltageRGA Spin Motor C Ph O.Skc VoltageYaw Attitude Error Voltage (O.Skc)Pitch Attitude Error Voltage (O.Skc)Roll Attitude Error Voltage (O.Skc)Descent Engine Arm Command (Pnl)Ascent Engine Arm Command (Pnl)Roll Attitude Cont Select (Pulsed)Pitch Attitude Cont Select (Pulsed)Yaw Attitude Cont Select (Pulsed)Roll Attitude Cont Select (Direct)

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Pitch Attitude Cont Select (Direct)Yaw Attitude Coat Select (Direct)Manual Thrust Command VoltageAuto Thrust Command VoltageDescent Engine On/Off To PropAscent Engine On/Off To PropX-Translation Command VoltageY-Translation Command VoltageZ-Translation Command Voltage .X-Translation OverrideJet Drivers Out 28V VoltageAuto/Manual Thrust VoltagePitch Gimbal Drive Actuator Position (O. kc)Roll Gimbal Drive Actuator Position (O.llkc)Sin IG Ix Res Out VoltageCos IG Ix Res Out VoltageSin MG Ix Res Out VoltageCos MG Ix Res Out Voltage

. Sin OG Ix Res Out VoltageCos OG Ix Res Out VoltageYaw Attitude Error To CESPitch Attitude Error To <JESRoll Attitude Error To CES± Delta Gamma From PGNS± Delta Theta From PGNS

- ± Delta- Phi From PGNS-CDU ZeroPGNS Pitch Attitude Error VoltagePGNS Yaw Attitude Error VoltagePGNS Roll Attitude Error VoltageLGC Altitude PositionLGC Altitude Velocity RateLGC Forward VelocityLGC Lateral VelocityDisplay Total, Sin Alpha (o.8kc) VoltageDisplay Total, Cos Alpha (O.Skc). VoltageDisplay Total, Sin Beta (O.Skc) VoltageDisplay Total, Cos Beta (O.Skc) VoltageDisplay Total, Sin Gamma (O.Skc) VoltageDisplay Total, Cos Gamma (O.Skc) VoltageAltitude To Display VoltageAltitude Rate To Display VoltageLatitude Velocity To Display Voltage+ AV To Display Voltage- AV To Display VoltageINSTRUMENTATION SUBSYSFrame Synch ID'sFormat IDCalibration 85$ High LevelCalibration 15$ High LevelCalibration 8556 Low LevelCalibration 15$ Low LevelGreenwich Mean TimeC & WE Master Alarm IDTCA's Failure

RADARSLdg Radar Range Data No GoodLdg Radar Velocity Data No GoodLdg Radar Heater CautionLdg Radar Velocity Xmtr RF Pwr Out Pwr

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PCPROPULSION SUBSYS - DESC STAGEHe Supply Tank Pressure.He. Supply Tank .Temperature . • ...He Primary Sol Valve CLOSED •He Secondary Sol Valve CLOSEDHe Regulator Outlet Manifold PressureHe Burst Disc RuptureFuel Tank No. 1'Liquid Low LevelFuel Tank.No. 2 Liquid Low LevelFuel Tank No. 1 Bulk TemperatureFuel Tank No. 2 Bulk-TemperatureFuel Control Valve Inlet Pressure

(XVI)2-91


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Oxid Tank No. 1 Liquid Low LevelOxid Tank No. 2 Liquid Low LevelOxid Tank No. 1 Bulk TemperatureOxid Tank No. 2 Bulk TemperatureOxid Control Valve Inlet PressureShut-Off Valves A/B Mid PositShut-Off Valves C/D Mid PositVariable Injector Actuator Posit

PCREACTION CONTROL SUBSYSHe Tank A PressureHe Tank B PressureHe Tank C PressureHe Tank D PressureHe Tank A TemperatureHe Tank B TemperatureHe Tank C TemperatureHe Tank D TemperatureHe Regulator A Cutlet PressureHe Regulator B Outlet PressureHe Regulator C Outlet PressureHe Regulator D Outlet PressureHe Shut-Off A-l ClosedHe Shut-Off B-l ClosedHe Shut-Off C-l ClosedHe Shut-Off D-l ClosedHe Shut-Off A-2 Not ClosedHe Shut-Off B-2 Not ClosedHe Shut-Off C-2 Not ClosedHe Shut-Off D-2 Not ClosedFuel Tank A PressureFuel Tank B PressureFuel Tank C PressureFuel Tank D PressureFuel Tank A TemperatureFuel Tank B TemperatureFuel Tank C TemperatureFuel Tank D TemperatureFuel Tank A QuantityFuel Tank B QuantityFuel Tank C QuantityFuel Tank D QuantityOxid Tank A PressureOxid Tank B PressureOxid Tank C PressureOxid Tank D PressureOxid Tank A TemperatureOxid Tank B TemperatureOxid Tank C TemperatureOxid Tank D TemperatureOxid Tank A QuantityOxid Tank B QuantityOxid Tank C QuantityOxid Tank D QuantityFuel Manifold A PressureFuel Manifold B PressureFuel Manifold C PressureFuel Manifold D PressureOxid Manifold A Pressure

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COMMUNICATIONS SUBSYSS-Band Steerable Antenna No TrackGimbal Pick-Off Output Y PositGimbal Pick-Off Output X PositSelected S-Band Rcvr AGC VoltageS-Band Rcvr AGC VoltageSelected S-Band-Xmtr RF PowerS-Band PA RF Power Output PowerST, PM, :ER., Selected S-Band Xpond Phase

PHAS

Parameter

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Structures

Thermodynamic

Electrical Power

Environmental Cont.-

Navigation & Guid.

Radars

S & C - CES

S & C - AGS

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Table 2.7-2

I I LAB OPERATIONAL MEASUREMENT SUMMARY

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Reaction Control

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0 1)2-95 96

Table 2.7-3

II LAB OPERATIONAL MEASUREMENTS SUMMARY

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(XVI)2-9T798J

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DNAL MEASUREMENT SUMMARY

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TAXI OPERATIONS

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(XVl)2-99/100

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(XVI)3-1

3. PRELAUNCH OPERATIONS

3.1 CHECKOUT. RATIONALE

The prelaunch checkout rationale for AES-LEM vehicles will be consistent with thepresent "Apollo Criteria" (Refs. U-U and k-5) and presently planned Apollo LEM pre-launch operations (Ref. U-6). The checkout'of-the AES-LEM vehicles'will begin inthe factory during subsystems installation and"continue through in-house vehicleacceptance testing. . At KSC, checkout will consist of subsystem, combined subsys-tem, and integrated system verification.

The major objectives of prelaunch checkout are to:

1. Verify system performance to assure accomplishment of mission objectives.

2. Establish compatibility of installed subsystems.

3. Demonstrate flight readiness.

U. Minimize the risk of:

a. Launching with a system which has degraded performance

b. "Holding" a. flight-ready vehicle

c. Missing the launch window.

The basic Manned Lunar Landing Program (Apollo-LEM) overlaps the Earth orbit andLunar AES flights so that the LEM prelaunch and mission operations are interspersedwith all the Phase I Labs; therefore the prelaunch checkout concepts and criteriawill follow from"the already approved plans for Apollo prelaunch operations. Theresulting AES prelaunch operations will minimize expenditures for facilities andGSE beyond those planned for the basic Apollo Program.

3.1.1 Concepts and Criteria -

The prelaunch operations are based- upon the following concepts and criteria:

a. End-to-end testing will be employed as the basic testing approach toprelaunch checkout. End-to-end testing is defined as applying aninput stimulus (or set of stimuli) to an identifiable functional flowpath or paths, within the system under test, and measuring the systemresponse at the end of that flow path or paths.

b. While end-to-end testing is the best approach for an overall prelaunch^checkout/' it will be necessary, in order to achieve all of the pre-launch checkout objectives, to deviate from the basic approach incertain cases. These cases will be justified on an individual basis.

c. Handling and transportation at all checkout sites will be minimized.

(XVI)3-2

d. Prelaunch checkout will be performed at as high an assembly level aspossible with the objective of maximum probability of malfunctiondetection. •' ="

e. The capability will exist to isolate a fault to a replaceable unitwhen a malfunction is detected. Exceptions must be justified on thebasis of significant reduction in weight or complexity of flightequipment.

f. Supporting subsystems will ..be tested to verify their supportingfunction. These supporting subsystems will be exercised over therequired range necessary to meet the checkout objectives during thecourse of other subsystem testing where these subsystems supply asupporting function.

g. In formulating checkout procedures and in the design of GSE, changingthe flight configuration for the purpose of checkout shall be mini-mized, i.e., GSE complexity is allowed for the sake of vehicle andflight hardware simplification.

h. Electrical, mechanical, and fluid connections of flight equipmentshall not be disconnected for equipment checkout after the vehiclehas been assembled and mated.

i. Spacecraft operation and performance of redundant elements withinreplaceable units will be verified.

j. Each branch of parallel functional processes will be verified inde-pendently.

k. All interfaces, both internal and external, will be verified.

1. All external equipment and ground complexes that interface-with thespacecraft will be verified prior to mating with the AES spacecraftor experiment.

m. Mission-essential GSE will be subjected to failure effects analysisto insure a vehicle-GSE fail-safe system.

n. Spacecraft telemetry (flight PCM) will be utilized during prelaunchcheckout to the extent possible with no ACE carry-on duplication,except for those checkout measurements that require an accuracy beyondthe capability of the flight PCM Subsystem.

o. Measurements required for prelaunch checkout will be available inreal-time to ACE.

p. Maximum utilization, consistent with safety, will be made of astro-nauts or equivalents and onboard controls and displays during check-out .

q. Maximum utilization will be made in the prelaunch checkout program ofall information and knowledge gained during the ground developmentprogram.

(XVI)3-3

r. Common checkout logic will be provided at all checkout locations forequivalent.levels of test.

s. System performance data obtained before and during prelaunch checkoutwill be maintained to develop a basis for trend analysis.

t. Data obtained during the factory checkout period will be applicableto KSC checkout.

u. Monitoring of critical parameters will be provided up to time oflaunch, so that malfunctions can be detected and remedial actiontaken.

^ ikumman-

(XVI)3-U

3.2 SUBSYSTEM CHECKOUT APPROACH , • '.

The development of a sound overall KSC prelaunch operations plan which effectivelyutilizes facilities and ground support equipment, must consider the four basic AES-LEM spacecraft requirements individually and collectively. Since the foundationfor facility and equipment requirements is dependent upon the AES-LEM subsystemconfigurations, this discussion considers a preliminary approach to the prelaunchcheckout of the AES-LEM subsystems. These approaches are based upon the RecommendedConfigurations of the Phase I Lab, Phase II Lab, Shelter, and Taxi which are pre-sented in the respective Design Analysis Reports, Vol'III through VI.

The approaches are later reflected in the development of individual AES-LEM space-craft flows (Paragraph 3«3)> overall checkout operations (Paragraph 3«^)> and.finally KSC facility and support equipment requirements (Paragraph 3«5)«

3.2.1 Propulsion/RCS Subsystems

3.2.1.1 Assumptions

Table 3-2-1 summarizes the vehicle/subsystem configurations assumed for checkout ofthe Propulsion/RCS Subsystems. Additional assumptions are:

a. Propulsion engine orifice sizing will be accomplished, and oxidizerto fuel (0/F) ratio checked, during cold-flow tests with substitutepropellents at the factory. This capability, although not requiredduring "normal" KSC checkout, will exist at KSC for contingencies(e.g., engine change).

b. The cryogenic system for helium storage (descent engine) does notrequire a full mission heat, leak check at KSC.

c. Wo static firing of the AES RCS or Propulsion Subsystems will beaccomplished at KSC. .

3.2.1.2 Objectives^

a. To verify pressure integrity of the RCS and Propulsion Subsystems(pressurization sections, propellant feed sections, and engine orthruster chamber assemblies).

b. To verify the functional capability of each component within eachsection.

c. To verify the operational capability of the engine or each thruster.

d. To verify supercritical helium tank insulation integrity.

3.2.1.3 Approach

a. Internal and external leakage will be determined on each component,line, tank, and disconnect. Leaks will be determined by volumetric-displacement, mass-spectrometer, and/or pressure-decay methods.

(xvi)3-5

b. Operation of each regulator, solenoid valve, relief valve, and checkvalve will be verified by flowing inert, conditioned gas through eachof these components. Solenoid valve cycling, however, will be keptto a minimum.

c. End-to-end inert gas flow checks, bypassing squib valves, will bemade through the propulsion pressurization, propellant feed, sectionsand engine (i.e., from upstream of regulator shut-off valves throughthe engine) to determine that no blockages exist in the system.

d. End-to-end checks of Guidance Navigation and Control Subsystem (GNCS)functions will include evaluation of command-signal response frompropulsion and RCS.

e. Because of the existence of positive expulsion bladders in the RCSpropellant tanks, end-to-end gas flow checks of the system are notfeasible. Therefore, the RCS propellant feed section/thrusterassemblies will be flow checked independently. Inert gas flow checkswill be made individually through each RCS thruster to verify injectorvalve response and ascertain that no blockages exist in the propellantflow path.

f. Cryogenic-tank insulation integrity will be verified by performing aheat leak check in a compressed time schedule.

g. Servicing of the propellant tanks will be accomplished at KSC for thefirst and only time (assuming no launch abort) on the launch pad.

h. Checkout of squib actuated vent, isolation, and initiator valves willbe accomplished during the checkout routine of other electro-explosivedevices (Paragraph 3.2.7).

3.2.2 Environmental Control Subsystems

3.2.2.1 Assumptions

Table 3«2-2 summarizes the vehicle/subsystem configurations assumed for checkout ofthe Environmental Control Subsystem (ECS).

3.2.2.2 Objectives

a= To verify the internal and external pressure integrity of each ECSsection (ARS, OSCP, WMS, and HTS) and each interface.

b0 To verify the support capability of those ECS sections which are usedin conjunction with checkout tests of other subsystems.

c. To demonstrate the functional performance of each section independentlyand interdependently.

(XVI)3-6

3.2.2.3 Approach

3.2.2.3.1 Atmosphere Revitalization Section (ARS).

a. Pressure distribution and circulation functional flow paths, compo-nents, and parts will be leak-checked and exercised to demonstratepressure and functional integrity.

b. Temperature control, and carbon dioxide and water removal functionswill be verified by simulating astronaut (or animal) metabolic inputsand outputs and establishing the ARS capability.

c. Operation of fans and pressure sensing circuitry will be checkedsimulating the range of differential and total pressures anticipatedfor the mission.

d. CM/EEM interfaces (or interdependencies) will be functionally verifiedand leak checked during docking tests. , . •

3.2.2.3.2 Oxygen Supply & Cabin Pressurization (OSCP).

a. The cabin, airlock, oxygen storage tanks and distribution lines willbe leak-checked to establish pressure, integrity of .hatch'seals., reliefvalves, check valves, dump valves, and quick disconnects.

b. Functional operation of over-pressure relief valves and dump valveswill be verified by determining cracking, open, and reseat pressuresand determining dump and repressurization times.

c. The OSCP will be exercised to demonstrate its ability to control thepressure and supply of oxygen to the ARS, .PISS, cabin, airlock, CM,and water tanks.

3.2.2.3.3 Water Management Section (WMS).

a. The pressure and functional integrity of the water functional flowpaths will be verified by circulating gas through the system toascertain that no blockages or leaks exist in the lines, water boilers,and associated components and WMS/HTS interfaces, and WMS/Fuel Cellinterface.

b. The ability of the water control module to achieve water regulationand process waste water will be verified.'

3.2.2.3.4 Heat Transport Section (HTS).

a. Before ECS provides cooling support for testing of other subsystems,the internal and external pressure and leakage integrity of .the HTSwill be established.

b. The ability of the HTS to supply coolant to other sections of ECS,EPS, radar, radiators and cold plates will be established by circulat-ing fluid through the system to ascertain that no blockages exist.

(XVl)3-7

c. Proper operation of backflow prevention, over-pressure relief valves,and coolant pump operation and control will be demonstrated.

3.2.3 Guidance Navigation & Control Subsystem (GNCS)

3.2.3.1 Assumptions

Table 3-2-3 summarizes the vehicle/subsystem configurations assumed for checkout ofthe GNCS. Additional assumptions are:

a. The Control Electronics Section (CES) and Command Control Section(CCS) will not require a section level checkout.

b. A complete checkout and verification of the CES and CCS will be per-formed during the integrated GNCS checkout where the required CESand CCS inputs will be available from other GNCS subsystems andsections.

c. The Abort Sensor Assembly (ASA) will undergo a complete laboratorybench check at KSC and will be re-installed in the vehicle prior tovehicle testing involving the Abort Guidance Section (ACS).

3.2.3.2 Objectives

3.2.3.2.1 Radar Section.

a. To verify radar transmitter and receiver operation in all modes.

b. To verify Landing Radar (LR) boresight accuracy.

c. To verify Rendezvous Radar (RR) pointing accuracy.

d. To verify Beacon Transponder operation.

3.2.3.2.2 Primary Guidance Navigation & Control Section (PGNCS) - GFE.

a. To functionally verify guidance, navigation and control operation.

b. To verify inertia! parameters.

c. To verify computer operational parameters.

d. To verify overall alignment between Inertia! Measurement Unit (IMU),the Alignment Optical Telescope (AOT) or Optical Tracker System (OTS).

3.2.3.2.3 Abort Guidance Section.

a. To verify the functional operation of the Abort Guidance Section (ACS)in all of its modes of operation.

b. To verify the various computer operational functions.

(XVI)3-8

c. To determine Abort Sensor Assembly (ASA) drift coefficients.

d. To verify the AGS operation under dynamic conditions.

3.2.3.2.14. Integrated GNCS.

a. To insure correct polarity of attitude references, control systems,and displays.

b. To verify the mechanical and electrical alignment throughout the GNCS.

c. To verify integrated GNCS operation.

d. To exercise the GNCS so that all RCS and propulsion'electrical inter-faces and control circuits are verified.

e. To verify Shelter Mission Programmer operation.

3.2.3.3 Approach

3.2.3.3.1 Radar Section.

a. The RR tracking circuitry will be verified with IF target generatorsignal inputs.

b. The boresight alignment of the LR transmitting and receiving antennaarrays will be checked to verify the electrical alignment of thebeams relative to each other and to a tilt mechanism.

c. RR pointing accuracy tests will be performed to verify the stringentbias and bias drift error limits of the RR.

d. Boresight shifts and tracking accuracies will be determined.

e. Compatibility between the RR and the CSM transponder will be verified.

f. The Shelter Beacon Transponder frequency, power output, sensitivityand delay time will be verified.

go Antenna hats will be employed in the verification of the RR and LRReceiver and Transmitter sections.

3.2.3.3.2 PGNCS (GFE).

a. LEM Guidance Computer (LGC) special test programs will be used in theverification of the individual functions and operations within thePGNCS.

b. LGC special test programs and optical tracker self-test will be usedin the verification of all modes of operation of the tracker.

c. Dynamic tracking tests will be performed to demonstrate the OTS sta-bility and tracking.

(xvi)3-9

d. Inertial parameters and computer operational parameters will "bedetermined to obtain a measure of performance and trend data.

e. IMJ and AOT or OTS alignment will be determined.

f. The PGNCS will be operated as a subsystem to verify all interfaces.

3.2.3.3.3 AGS.

a. An AGS functional test will be performed to verify proper operationin each mode. In the Inertial Reference Mode a special test programin the Abort Electronics Assembly (AEA) memory will verify all AESoutput interfaces and proper mode operation.

b. The AEA self-test programs will be used to verify AEA operation,whenever the computer is required in any test sequence.

c. ASA drift coefficients will be verified and determined, with the ASAin the AES-LEM, so that these coefficients can be loaded into the AEAfor drift compensation.

d. The AES-LEM will be positioned by a three-axis positioner so that theASA accelerometer outputs can be checked and accelerometer errorsources determined.

e. While in the positioner, significant changes in ASA gyro scale factor,bias and misalignment coefficients will be determined. Gyro scalefactor, per se, can not be determined while the ASA is installed inthe AES-LEM vehicle.

3.2.3»3.^ Integrated GNCS.

a. GNCS polarity, descent engine gimbaling, PGNCS automatic translationcommands and displays will be verified with vehicle inputs from thepolarity test fixture.

b. Alignment of the RR to the navigation base, LR antenna and descentengine will be verified.

c. End-to-end checkout of the GNCS will"be accomplished in all modes ofoperation through radar inputs into the LGC, special computer testprograms and the resulting LGC and AEA inputs to the CES.

d. The Program Coupler Assembly operation will be verified by monitoringGNCS responses to uplink commands to the command receiver of theShelter.

e. The X-Y Scanner operation will be verified as described in 3.2.5.3d

f. The Sequencer Assembly operation will be verified by "initializing"the scanner and monitoring the taxi status bus.

(XVI)3-10

3.2.4 Electrical Power Subsystems

3.2.4.1 Assumptions

Table 3.2-4 summarizes the vehicle/subsystem configurations assumed for checkout •of the Electrical Power Subsystems (EPS). Additional assumptions are:

a. The Power Distribution Section (PDS) serves a supporting function andafter initial verification, will be checked in conjunction with othervehicle checks.

b. Development of the cryogenic storage system for fuel-cell reactantsdoes not require a full mission heat leak check at KSC.

3.2.4.2 Objectives

a. To monitor and verify PDS operation while functioning under subsystemloads.

b. To verify the compatibility of the PDS and Power Generation Section(PGS) under load and no-load conditions.

c. To verify the continuity and isolation of the wiring harness and allexternal interfaces.

d. To verify the activation of the flight batteries,

e. To verify the internal and external pressure integrity of the PGS,i.e., the fuel-cell and reactant storage subsections.

f. To verify the functional operation of the electrical and fluid controland monitoring devices of the fuel cells.

g. To verify cryogenic storage tank insulation integrity,

h. To verify operating performance of the fuel-cells.

3.2.4.3 Approach

a. The malfunction detection and correction, and switch-over functionsof Electrical Control Assemblies (EGA) will be verified.

b. Inverter operations will be checked and monitored.

c. Continuity and isolation of wiring harnesses and external interfaceswill be established.

d. After activation, batteries will be partially discharged to verifyproper activation.

e. Battery charger operation will be functionally verified.

(xvi)3-il

f. Internal and external leak checks (with inert gas) will he made onthe fuel-cell and reactant storage tanks, components, distrihutionnetworks and interfaces to establish pressure integrity.

g. Components (flow control valves and regulators, purge valves,pressure switches, etc.) throughout functional flow paths will beexercised to establish functional integrity of each fuel-cell andstorage subsection individually.

h. Servicing functions and cryogenic-tank insulation integrity will beverified by performing a heat leak test in a compressed time schedule.

i. The operation of the fuel cells will be verified by generatingelectrical power to subsystems loads, simulating critical powerdistribution, redundant, and emergency modes of electrical controllogic.

3.2.5 Communications Subsystem

3.2.5.1 Assumptions

Table 3-2-5 summarizes the vehicle/subsystem configurations assumed for checkoutof the communications subsystem.

It is assumed that the S-Band steerable antenna and its related equipment is. on allPhase I and Phase II Labs; however, the inclusion of the S-Band equipment is depend-ent on the communications circuit margin requirements for different missions of thePhase I and Phase II Labs.

3.2.5.2 Objectives

a. To verify that the Communication Subsystem is functioning properly inall modes of operation.

b. To verify S-Band steerable antenna operation.

c. To verify command receiver and up-data link operation.

d. To demonstrate communications compatibility with Manned Space FlightNetwork (MSFN).

e. To verify the Electro-Magnetic Compatibility (EMC) of the AES-LEMvehicle.

3.2.5.3 Approach

a. The proper functioning of the Communication Subsystem will be verifiedend-to-end by operating the subsystem in all its modes. Antennaadapters will be placed over the S-Band steerable antenna, the twoS-Band inflight and the two VHP inflight antennae. The LEM erectableantenna bulkhead connector and the S-Band antenna adapters will becoupled via transmission lines to an RF switch and then to thecommunications console. The operation functions will be performedutilizing a selected S-Band and VHP antenna.

(XVI)3-12

b. A complete end-to-end test will be performed with the ground stationto demonstrate compatibility between the vehicle Communication Systemand the MSFN.

c. The S-Band steerable antenna's dynamic tracking characteristics willbe tested with the use of an S-Band source located at a distance fromthe vehicle and by moving the vehicle within the antenna's trackinglimits.

d. The Shelter's X-Y Scanner will be functionally verified. With the S-Band steerable antenna locked on to an S-Band source the vehicle willbe moved so that "lock on" is broken. An up-link command will then besent initiating the X-Y Scanner operation.

e. Vehicle EMC will be verified by operating all equipment includingradiating devices and monitoring system performance.

3.2.6 Instrumentation Subsystems

3.2.6.1 Assumptions

Table 3«2-6 summarizes the Vehicle/Subsystem configurations assumed for checkout ofthe Instrumentation Subsystem. Additional assumptions are:

a. The Instrumentation Subsystem serves a supporting function and, afterinitial verification, will be checked in conjunction with other vehiclechecks.

. b. The instrumentation for the Phase I Labs will be recalibrated at KSCafter the manufacturing modification.

3.2.6.2 Objectives

a. To verify that the Operational Instrumentation is ready to supportspecific vehicle checkout.

b. To verify that all instrumentation inputs, outputs and all redundantpaths are functioning.

c. To verify the proper functioning of the Caution and Warning Section,Pulse Code Modulation & Timing Electronics Assembly (PCMTEA), andVoice Storage Recorder (VSR).

d. To verify Experiment-Instrumentation section.

3.2.6.3 Approach

a. All Caution and Warning (C&W) alarms will be verified by applying an"alarm enable" to each C&W channel while inhibiting all other channels.

b. The PCMTEA will be verified by monitoring its downlink and GSE data.

(xvi)3-i3

c. The Experiment Instrumentation for the Phase I and Phase II Labs willbe verified by the application of stimuli and monitoring the down-link.

d. The VSR will be checked by monitoring read-head output after applyinga time limited test signal to the recording head.

3.2.7 Structural/Mechanical and Electro-Explosive Devices Subsystems

3.2.7.1 Assumptions

Table 3-2-7 summarizes the vehicle/subsystem configurations assumed for checkout ofthe'Structural/Mechanical and Electro-Explosive Devices Subsystem.

3.2.7.2 Objectives

a. To verify the mechanical connections between AES-LEM stages andbetween the vehicle and the Spacecraft-LEM Adapter (SLA).

b. To establish leakage integrity across soft and hard dock seals,hatches, tunnels, and airlocks.

c. To verify the mechanical and electrical docking interfaces and hatchoperations.

: d. To verify the proper functioning of the thermal covers for hatches,windows, and RCS clusters.

e. To verify mechanical functioning of airlocks.

f. To demonstrate landing gear deployment.

g. To demonstrate performance of each type of electro-explosive device,

h. To verify the firing circuits for each electro-explosive device.

3.2.7.3 Approach

a. Electrical and mechanical interfaces between AES-LEM and GSM will beverified and leakage integrity established during hard docking betweenthe vehicles or stages involved.

b. Fit checks between SLA and descent stage will be made to establishmechanical compatibility before stacking the AES-LEM and CSM withinthe SLA.

c. Performance of electro-explosive devices (BED) will be demonstratedin "destructive tests" using statistical sampling methods prior to

installation in the vehicle.

(XVI)3-llti

3-2.8 Crew Provisions Subsystems

3.2.8.1 Assumptions

Table 3*2-8 summarizes the vehicle/subsystem configurations assumed for checkout ofthe Crew Provisions Subsystem. It is assumed that all GFE items will be checked outas components before delivery to GAEC or at field installations.

3.2.8.2 Objectives

a. To verify the functional capability of the pressurized compartmentsincluding installation of equipment and station arrangement; instru-ment panels and consoles; lighting fixtures, illumination and color;and interstage circuitry.

b. To verify the functional interfaces between the Portable Life SupportSystem (PLSS) and AES LEM.

c. To verify the compatibility and functional interfaces between the AESLEM and soft and hard suits.

d. To verify the component and installation integrity of harnesses,belts, slings, seats, hand/foot holds, and other Crew Provisionequipment.

e. To verify the leakage integrity and functional operation of the WasteManagement Section.

f. To verify the functional operation of Special Devices (e.g., waterprobe).

g. To assure completeness of quantities of equipment and mechanicalinterface provisions.

3.2.8.3 Approach

a. All lighting will be inspected to verify lamp function, intensity ofillumination, color, and coordination of dimming control.

b. Pressure checks at operating and higher pressures will be made at theliquid and gas interfaces between the PLSS and vehicle, and PLSS andsuit.

c. Pull tests will be performed on all crew harnesses, belts, slings,and foldable bunk components.

d. Components and plumbing of the Waste Management Section will beleak checked with an inert gas and functionally checked with testfluids.

e. Water probe operation in the preparation of food, drinking, and PLSSrecharge will be checked.

f. Space suit and PLSS stowage provisions and vehicle interfaces will bechecked.

(XVI)3-15

3.2.9 Experiment Interfaces

As related to checkout functions, the AES-LEM experiment categories are defined asfollows:

a. Independent experiments are those which depend on the AES-LEM vehiclesolely for storage and/or mounting. An instrumentation interface withthe vehicle may or may not exist.

"b. Dependent experiments are those which depend on the vehicle forpower, environmental control, orientation, and/or other supportingfunctions in addition to storage, mounting, and instrumentation.

c. Integrated experiments are those which are essentially interrelated(either in functional flow path or structurally) within the AES-LEMvehicle subsystems, such that the experiment cannot be conductedwithout vehicle function.

3.2.9.1 Assumptions

Table 3«2-9a summarizes the experiment interface checkout assumptions as applicableto each category of experiments.

3.2.9.2 Objectives

Table 3«2-9b summarizes the experiment interface checkout objectives as applicableto each category of experiments.

3.2.9.3 Approach

Table 3«2-9c summarizes the vehicle/experiment checkout approach as applicable toeach category of experiments.

(XVI)3-16

TABLE 3.2-1

PROPULSION/RCS CHECKOUT CONFIGURATIONS

LEM Ascent Stage Propulsion

LEM RCS

LEM RCS (1/2 PropellantCapacity and Thrusters)

LEM RCS (Double PropellantCapacity)

RCS Vent Valves

ISM Descent Stage Propulsionwith Vent Valves

Phase I Lab

X

Phase II Lab

X-

Shelter

X

X

X

Taxi

X

X

X

NOTE: Based on Vehicle Recommended Configuration; See Vols . Ill, IV, V & VI.

(XVI)3-17

TABLE 3.2-2

ECS CHECKOUT CONFIGURATIONS

ATMOSPHERE REVITALIZATION SECTION

Suit Circuit Assy

Cabin Recirculation Assy

CSM/LEM Recirculation Duct

02 SUPPLY & CABIN PRESSURIZATIOWSECTION

LEM System

Additional LEM GOX Tanks

Airlock

'Two -Gas System

GOX from Fuel Cell Storage

WATER MANAGEMENT SYSTEM

LEM System

Additional Potable Water Tanks

Modified CSM System

HEAT TRANSPORT SECTION

ASA Bypass

Radiators

Separate Fuel Cell Cooling Loop

RTG Heat Pipe

Additional Cold Plates

Phase I Lab

X

X

X

X

X

X

X

Phase II Lab

X

X

X

X

X

X

X

X

X

X

Shelter

X

X

X

X

X

X

X

X

Taxi

X

X

X

X

X

X

NOTE: Based on Vehicle Recommended Configuration; See Vols. Ill, IV, V & VI.

j/UUTWlCUl.

(XVI)3-18

TABLE 3.2-3

GN&C CHECKOUT CONFIGURATIONS

• RADAR SECTION

Rendezvous Radar

Landing Radar

Beacon Transponder

• PRIMARY GN&C (GFE)

Inert ial Measurement Unit

LEM Guidance Computer

Power and Servo Assembly

Pulse Torque Assembly

Display and Keyboard

Alignment Optical Telescope

Optical Tracker

Navigation Base

• ABORT GUIDANCE SECTION

Abort Sensor Assembly

Abort Electronics Assembly

Data Entry and Display Assy

• CONTROL ELECTRONICS SECTION

Attitude and TranslationControl Assembly

Rate Gyro Assembly

Ascent Engine Latching Device

Descent Engine Control Assy

Gimbal Drive Actuator

Attitude Control Assembly

Thrust Translation Control Assy

• COMMAND CONTROL SECTION

Program Coupler Assy

X-Y Scanner Assy

Sequencer Assy

Phase I Lab

X

X

X

X

X

X

X

X

X

Phase II Lab

X

X

X

X

X

X

X

X

X

Shelter

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

Taxi

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X


(xvi)3-l9

TABLE 3.2-U

EPS CHECKOUT CONFIGURATIONS

All-Battery PGS

Fuel Cells with CryogenicStorage of Reactants _

Fael Cells with AmbientStorage of Reactants

Radioisotope Thermal GeneratorCRTG)

Phase I Lab

X

Phase II .Lab

X

Shelter

X

X

Taxi

X

X


(XVI)3-20

TABLE 3.2-5

COMMUNICATIONS CHECKOUT CONFIGURATIONS

S-Band

VHP

Signal Processor Assy

TV Camera

X-Y Scanner *

Command Receiver

Up -Data Link

AES-LEM/CSM Intercom

S~Band Erectable Antenna

Phase I Lab

X

X

X

X

Phase II Lab

X

X

X

X

Shelter

X

X

.X

X

X

X

X

Taxi

X

X

X

X

X

NOTE: Based on Vehicle Recommended Configuration; See Vols . Ill, IV, V & VI.,

* Listed for reference only; part of GNCS.

(XVI)3-21

TABLE 3.2-6

INSTRUMENTATION CHECKOUT CONFIGURATIONS

Pulse Code Modulation &Timing Electronic Assy

Signal Conditioning ElectronicsAssy

Caution & Warning ElectronicsAssy

Voice Storage Recorder

Experiment Pulse Code Modu-lation & Timing ElectronicAssy

Experiment Data StorageEquipment

Phase I Lab

X

X

X

X

X

X

Phase II Lab

X

X

X

X

X

X

Shelter

X

X

X

X

Taxi

X

X

X

X


_

(XV1)3-22

TABLE 3.2-7

STRUCTURAL/MECHANICAL AND EEC CONFIGURATIONS

STRUCTURAL

Window Covers

Hatch Covers

MECHANICAL

Interstage Exp. Bolts

Guillotine Cable Cutter

Landing Gear

Propulsion Pyro

Components

Phase I Lab

X

Phase II Lab

X

Shelter

X

X

X

X

Taxi

X

X

/

X

X

X

X

NOTE: Based on Vehicle Recommended Configuration; See Vols . Ill, IV, V & VI.

(xvi)3-23

TABLE 3.2-8

CREW PROVISIONS CHECKOUT CONFIGURATIONS

•

Modified Ext & Int Lighting

Waste Management System

Flexible, Stowable Bunks

Hard Suit (GPE)

Extra PLSS (GFE)*

PLSS LiOH Cartridges (GFE)*

PLSS Batteries

Crew Support & Restraint

Work Tops

Pressurized Compartments- Equipt Ins t Is

Medical Facilities (GFE)**

Mod Instr Panel & Consoles ***

Ingress/Egress Provisions(Airlock, etc)

Food (GFE)*Extra Constant-Wear Garments

(CWG; GFE)

Phase I Lab

X

X

X

X

X

X

X

X

X

Phase II Lab

X

•

X

X

X

X

X

X

X

X

X

Shelter

X

X

X

X

X

X

X

X

X

X

X

' X

X

X

Taxi

X

X

X

X

X

X

* Equipt presently in LEM, but different quantities.

''=** Anticipated variation from LEM.

*** Result of Control & Display changes.

NOTE: Based on Vehicle Recommended Configurations; See Vols. Ill, IV, V & VI

(XVI)3-2U

TABLE 3-2-9a

EXPERIMENT INTERFACE ASSUMPTIONS & CATEGORIES

ASSUMPTIONSEXPERIMENT CATEGORY

Independent Dependent Integrated

If electrical power is provided

to an experiment only for instru-

mentation control or stimuli, it

is considered to be an instrumen-

tation interface for the purpose

of experiment categorization.

The experimenter will checkout

his own experiment package, using

bench test equipment, to the de-

gree possible in an earth en-

vironment, prior to installation

in the vehicle„

EMI testing will be accomplished

for each mission-vehicle/experi-

ment package at the system test

level.

Experiments requiring active

thermal control will be supported

by the ECS thermal control loop.

When total experiment checkout

cannot be accomplished, com-

ponent checkout within the experi-

ment package will be conducted

prior to installation and an inte-

grated checkout will be performed

after installation.

X

X

X X

X X

X

(XVl)3-25

TABLE 3-2-9b

EXPERIMENT INTERFACE CHECKOUT OBJECTIVES & CATEGORIES

OBJECTIVESEXPERIMENT CATEGORY

Independent Dependent Integrated

To demonstrate that experimentcan be stored, set up and operat-ed in the AES vehicle.

To verify the data functionalflow paths vhen an instrumenta-tion interface with the vehicledata recording and transmissionequipment exists.

To verify that no EMI problemsexist after installation of ex-periments in the AES flight ar-ticle.

To verify that the mechanical in-terface points required for stor-age and operation of the experi-ment are compatible with the de-livered experiment package.

To verify the operational com-patibility between supportingsubsystem(s) and experiment(s).

To verify the integrity of thefluid flow paths when a directfluids interface exists betweenexperiment(s) and subsystem(s).

To functionally demonstrate ex-periment ( s)•deployment mechanism.

To verify the capability of thevehicle to provide support to theexperiment„

To verify that the crew/experi-ment/vehicle combination oper-ates properly in all modes ofoperation, to the extent possi-ble in Earth environment.

X

X

X

X

X

X X

X

(XVI)3-26

TABLE 3.2-9c

EXPERIMENT INTERFACE APPROACH & CATEGORIES

APPROACHIndependent

EXPERIMENT CATEGORY

Dependent Integrated

Storage of an experiment in theAES-LEM vehicle will be verifiedby actual installation. Since aprior check will have been madeusing an experiment mock-up orprototype, the final fit checkdemonstrates that the deliveredexperiment package is compatiblewith the flight article„

To verify the capability of theLab to accept an experiment con-figured for operation, the exper-iment will be set up and operated,if possible, in earth environment.

After experiment installation, ascreen room test will be performedwith all subsystems and experi-ments operating to verify that noEMI problems exist.

Instrumentation data flow pathswill be end-to-end checked duringexperiment operation. If the ex-periment is incapable of Earthenvironment operation, artificialstimuli will be used, if possible,to accomplish the end-to-endcheck.

If special fluid systems or loopsexist to support a particular ex-periment, these loops will alsobe checked out independently. Acomparison of system output param-eters versus experiment inputrequirements will be made. Inaddition, leak checks and compo'-nent functional tests will beperformed.

X

X X X

X X

(XVl)3-27

TABLE 3.2-9c

(Cont'd)

APPROACH

• The supporting subsystems shall beoperated in conjunction with theexperiment to demonstrate that thesubsystem/experiment combinationcan function normally in theflight vehicle.

• Mechanical deployment devices willbe exercised/ if possible, todemonstrate their proper operation.

• When the nature of the experimentprecludes its complete operationoutside of the vehicle, an inte-grated mission-oriented test afterinstallation will be performed,within the constraints of Earthenvironment, exercising the partic-ular vehicle/experiment combination.

EXPERIMENT CATEGORY

Independent Dependent

X

X

Integrated

X

X

X

j/uunman.

(XVI)3-28

3.3 CHECKOUT OPERATIONS

The checkout flow through the KSC facilities for each of the basic AES-LEM space-craft/launch vehicle configurations is shown in Fig. 3-3-1 and discussed in theparagraph to follow.

3.3.1 Assumptions

The checkout flow for each type of AES-LEM through the KSC facilities is based onthe checkout approach for the individual subsystems (Sec. 3.2) and reflects max-imum utilization of currently planned KSC facilities and equipment. The followingassumptions are made relative to the operations in each of the facilities:

3.3.1.1 General

a. The Phase I Labs have not been cycled through any part of the KSC pre-launch checkout flow before the modification period.

b. As a goal, all vehicle/experiment comprehensive testing -will be accom-plished in the MSOB. However, it is recognized that specific exceptions,may require other facilities (e.g. HTB, KFTF).

c. All experiments are installed before integrated testing.

. d. Storage of the descent stage will not require a facility normally usedfor checkout purposes.

3.3.1.2 Hypergolic Test Building (HTB)

a. The checkout of the Phase II Lab RCS requires two more days than thePhase I Lab because of the multiple tank configuration of the propellantfeed section in the former.

b. The duration of concurrent checkout for the Shelter RCS and descentpropulsion is established by the descent stage checkout.

c. The duration of checkout for the propulsion subsystems of the Taxi isestablished by the sequential checkout of the ascent engine and RCS.

3.3.1.3 Manned Spacecraft Operations Building (MSOB)

a. Modification of the Phase I Labs at KSC necessitates a Phase I Labintegrated subsystems compatibility verification. As a result, thetotal Phase I Lab MSOB checkout will be 7 days longer than the PhaseII Lab MSOB checkout.

b. Phase II Labs and Shelters require two additional days in the AltitudeChamber for fuel cell checkout.

(XVl)3-29

c. Top deck buildup for the spacecraft launched on Saturn V will be accom-plished in the MSOB.

3.3.1. Fuel Cell Systems Test Facility (FCSTF)

a. Fuel cell, reactant storage, and ECS configuration of the Phase II Laband Shelter requires that the ascent and descent stages be mated priorto checkout in the FCSTF.

b. The cryogenic storage configuration of the Phase II Labs requires fourmore days of checkout, time than the ambient storage configuration ofthe Shelter because of the insulation integrity checks required of theformer.

3.3.1.5 RF Systems Test Facility (RFSTF)

The additional time required by the Taxi at the RFSTF is for checkout of the Rendez-Radar.

3.3.1.6 Launch Complex

a. The checkout times shown for the VAB, Launch Pads 39 and 37 are governedby Saturn launch vehicle checkout requirements. The spacecraft checkout(CSM/LEM) is done concurrently with booster operations.

b. The difference in times between Launch Complex 37 and 39 operations isattributed to launch vehicle activities.

3.3-2 Phase I Lab (Fig. 3-3-1A)

3.3.2.1 Ascent Stage

3-3.2.1.1 Hypergolic Test Building (West Cell). Upon completion of the KSC modi-fication, the ascent stage will be transported to the West Cell of the HTB. Check-out and servicing of the ECS Heat Transport Section will be performed to providecoolant capability during subsequent tests. Leak and functional checks will beperformed on the RCS, Oxygen Supply and Cabin Pressurization Section and Water Man-agement Section plumbing. The ascent stage will then be transported to the MSOB.

3.3.2.1.2 Manned Spacecraft Operations Building. After the ascent stage is trans-ported to the MSOB it will be installed in the cleaning fixture, inverted and thenmoved to the docking fixture, for the Lab-CSM docking test. Upon completion of thedocking test the ascent stage will be moved to the cleaning fixture for reinversionand Water Management, Heat Transport and Oxygen Supply and Cabin PressurizationSections checkout. The ascent stage will then be prepared for mating with the de-scent stage.

3.3.2.2 Descent Stage

3.3.2.2.1 Storage Area. After completion of the KSC modification period the Labdescent stage will be stored until required for SLA fit check in the MSOB. No spe-cific area has been designated for storage.

(XVI)3-30

3.3.2.2.2 Manned Spacecraft Operations Building. Upon arrival of the descentstage at the MSOB it will "be transported to the SLA stand for a descent stage/SLAfit check. The stage will then be moved to the proper work stand for mating withthe ascent stage. The Heat Transport and Water Management Sections of the ECS will"be checked prior to mating the descent and ascent stages.

3.3.2.3 Mated Lab

3.3-2.3.1 Manned Spacecraft Operations Building. After mating, the Heat TransportSection will be leak checked and serviced. The Power Distribution Section and In-strumentation Subsystem capability to support vehicle testing will be verified.Checks will be made to verify Alignment Optical Telescope calibration and the align-ment of the S-Band antenna and experiment equipment. Electrical and electronicsubsystem tests, GWCS polarity, and experiment verification will be performed.

The Phase I Labs will undergo a complete vehicle system integrated checkout. Allvehicle electronic subsystems will be operated in an integrated fashion so that allpossible system modes are verified. A "plugs out" vehicle EMC test will be per-formed in a shielded chamber. To demonstrate the ability of the electrical andelectronic subsystems to work together as a system, a mission-oriented "plugs out"test will be performed. The AES Lab will then be installed in the altitude chamber.

After installation in the altitude chamber, ECS, cabin leak, and hatch operationtests will be performed prior to evacuating the chamber. The chamber will be evac-uated with the Lab unmanned, and ECS functional tests will be performed to demon-strate the capability of the ECS to support manned altitude tests. The structuralintegrity of the Lab, during emergency re-pressurization of the chamber, will bedemonstrated in the unmanned altitude run. The manned altitude tests will demon-strate ECS and Crew Provisions capabilities with a man in.the loop. After themanned altitude tests the vehicle will be deserviced and transported to the RFSTF.

3.3.2.3.2 RF System Test Facility. The vehicle will be mounted on the RFSTF three-axis positioner. A communication test with the MSFN will be performed to verify thecompatibility between the -Lab Communication Subsystem and MSFN. The S-band steer-able antenna will be functionally tested, utilizing an S-band source located atsome distance from the vehicle, and motion inputs to the positioner. An AGS dynam-ic test will be performed with the' use of the three-axis positioner. The vehiclewill be oriented to various positions so that a gross determination of ASA scalefactors can be made. Lab EMC and Lab/CSM EMC tests will be performed at the RFSTF.Upon completion of testing at the.RFSTF, the vehicle will be transported to theMSOB.

3-3-2.3.3 Manned Spacecraft Operations Building. The Lab will be installed on aworkstand in the MSOB. Pyrotechnic circuitry will be verified and the explosivedevices (minus initiators) mechanically installed. After thermal shield installa-tion and exterior refurbishment, the Lab will be mated with the CSM/SLA in theSLA stand. For Saturn IB missions, the mated spacecraft (Lab/SLA/Boilerplate CSM)will be transported to LC-3T- For Saturn V missions the spacecraft (Lab/SLA/CSM)will be transported to the VAB.

3.3.2.3. Vehicle Assembly Building (Saturn V). The Spacecraft (S/C) will bestacked on the Launch Vehicle (LV) which will be in place on the Mobile Launcher(ML) in a high-bay area of the VAB. While integrated tests of the CSM and LV are

(xvi)3-3l

being performed, checks will be made on the Lab electronic subsystems, and a pres-sure decay test made on the RCS. EMC checks will be performed on the Lab duringthe CSM/LV "plugs out" test. The Launch Escape System (LES) will be installed andML swing arm tests will be performed. The space vehicle will then be moved to thelaunch pad.

3.3.2.3.5 Launch Pad 39 (Saturn V). The space vehicle will be transported to thelaunch pad on the ML and placed in position. Range/Space Vehicle RFI tests will beperformed before the Mobile Service Structure (MSS) is moved into position. Afterthe MSS is positioned, a checkout of the Lab subsystems will be performed. Lab RFsubsystem checks will be performed, using antenna hats and SLA reradiating antennas.Following subsystems verification, the Lab will participate in a flight readinesstest and then hypergolic servicing will be accomplished. A countdown demonstrationwill be performed to verify the capability of accomplishing terminal countdown pro-cedures in the allotted time. EPS flight batteries will be installed. After afinal subsystems checkout, all ACE carry-on equipment will be removed from the SLA.Lab pyrotechnic initiators and pyrotechnic batteries will then be installed. Finalfluid servicing will be accomplished, and the Lab cabin will be closed out and pres-surized. All GSE work platforms will be removed from the SLA and the SLA will beclosed out.

3.3.2.3.6 Launch Pad 3TB (Saturn IB). The Spacecraft (S/C) will be stacked on theLaunch Vehicle (LV) at Launch Pad 37B. The Launch Escape System (LES) will be in-stalled and launch tower swing arm tests will be performed. While integrated testsof the CSM and the LV are being performed, the Lab subsystems will be checked. Theservice structure will be removed and Range/Spacecraft RFI tests and "plugs out"tests performed. The service structure will be repositioned and pressure decaytests made on the.Lab RCS. A countdown demonstration will be performed to verifythe capability of accomplishing terminal countdown procedures in the allotted time.The Lab. will then participate in the Space Vehicle Simulated Flight. All hypergolicservicing will then be accomplished and EPS flight batteries installed. After afinal subsystems checkout, all ACE carry-on equipment will be removed from the SLA.Pyrotechnic initiators and batteries will then be installed. Final fluid servicingwill be accomplished, and the Lab cabin will be closed-out and pressurized. AllGSE and work platforms will be removed from the SLA and the SLA will be closed-out.

3.3.3. Phase II Lab (Fig. 3-3-1B)


3.3.3.1.1 Hypergolic Test Building (West Cell). Upon arrival at KSC, the ascentstage will be transported to the West Cell of the HTB where the receiving inspec-tion of the stage will be performed. Checkout and servicing of the ECS Heat Trans-port Section will be performed to provide coolant capability during subsequent tests.Leak and functional checks will be performed on the RCS, Oxygen Supply and CabinPressurization Section, and Water Management Section plumbing. The ascent stagewill then be transported to the MSOB.

3.3.3.1.2 Manned Spacecraft Operations Building. After the ascent stage is trans-ported to the MSOB it will be installed in the cleaning fixture, inverted and thenmoved to the docking fixture for the Lab-CSM docking test. Upon completion of thedocking test, the ascent stage will be moved to the cleaning fixture for inversionand Water Management, Heat Transport, and Oxygen Supply and Cabin PressurizationSections checkout. The ascent stage will then be prepared for mating with thedescent stage.

s ,'Aumman.

(XVI)3-32

3.3-3-2 Descent Stage

3.3.3.2.! Storage Area. Upon arrival at KSC, the Lab descent stage will be trans-ported to a storage area where the receiving inspection of the stage will be per-formed. The Lab descent stage will be stored until required for SLA fit check inthe MSOB. No specific area has been designated for storage.

3.3.3.2.2 Manned Spacecraft Operations Building. Upon arrival of the descent stageat the MSOB it will be transported to the SLA stand for a descent stage/SLA fitcheck: The stage will then be moved to the proper workstand for mating with theascent stage. The Heat Transport and Water Management Sections of the ECS will bechecked prior to mating the descent and ascent stages.

3.3.3.3 Mated Lab

3.3.3.3.1 Manned Spacecraft Operations Building. The ascent stage and descentstage will "be mated and transported to the FCSTF.

3.3.3.3.2 Fuel Cell System Test Facility. The Lab will be installed on a checkoutstand. Electrical functional checks will be performed on the reactant storage sec-tion, including the associated controls and displays. After leakage and mechanicalfunctional checks are performed, the storage section will be serviced with cryogenicfluids and heat leak test performed to establish insulation integrity. Flow checkswill be performed without start up of the fuel cells. The vehicle will then be de-serviced and transported to the MSOB.

3.3.3.3.3 Manned Spacecraft Operations Building. Upon arrival at the MSOB the Labwill be installed on a checkout stand, and the Heat Transport Section will be leakchecked and serviced. The Power Distribution Section and Instrumentation Subsystemcapability to support vehicle testing will be verified. Checks will be made toverify Alignment Optical Telescope calibration, and S-Band antenna and experimentalignments. Electrical and Electronic Subsystem tests, GNCS polarity, and experi-ment verification will be performed.

An integrated GNCS test will be performed wherein all functional flow paths andmodes of operation are verified. To demonstrate the ability of the electrical andelectronic subsystems to work together as a system, a mission oriented "plugs out"test will be performed. The Lab will then be installed in the altitude chamber.

After installation in the altitude chamber, ECS, cabin, airlock, and hatch leakageand functional tests will be performed prior to evacuating the chamber. The cham-ber will be evacuated with the Lab unmanned, and ECS functional tests will be per-formed to demonstrate 'the capability of the ECS to support manned altitude tests.The structural integrity of the Lab, during emergency re-pressurization of thechamber, will be demonstrated in the unmanned altitude run. The manned altitudetests will demonstrate ECS and Crew Provisions capabilities with a man in the loop.The fuel cells will be operated (with no cryogenic servicing involved) in the alti-tude chamber concurrently with ECS checkout. Electrical power will be generatedto other subsystems simulating critical power distribution requirements, redundantand emergency modes of electrical and control logic. After the manned altitudetests, the vehicle will be deserviced and transported to the RFSTF.

3.3.3»3.^ RF Systems Test Facility. .The vehicle will be mounted on the RFSTF three-axis positioner. A communication test with the MSFN will be performed to verify

(xvi)3-33

the compatibility between the Lab Communication Subsystem and MSFN. The S-bandsteerable antenna will be functionally tested, utilizing an S-band source locatedat some distance from the vehicle and motion inputs to the positioner. An ACSdynamic test will be performed with the use of the three-axis positioner. Thevehicle will be oriented to various positions so that a gross determination of ASAscale factors can be made. Lab EMC and Lab/CSM EMC tests will be performed at theRFSTF. Upon completion of testing at the RFSTF, the vehicle will be transported tothe MSOB.

3-3-3-3-5 Manned Spacecraft Operations Building. The Lab will be installed on awork stand in the MSOB. The pyrotechnic circuitry will be verified and the explo-sive devices (minus initiators) mechanically installed. After thermal shield in-stallation, radiator servicing, and exterior refurbishment, the Lab will be matedwith the CSM/SLA in the SLA stand. For Saturn IB missions, the mated spacecraft(Lab/SLA/Boilerplate CSM) will be transported to LC-37- For Saturn V missions thespacecraft (Lab/SLA/CSM) will be transported to the VAB.

3.3.3.3.6 Vehicle Assembly Building (Saturn V). The Spacecraft (s/C)will bestacked on the Launch Vehicle (LV) which will be in place on the Mobile Launcher(ML) in a high-bay area of the VAB. While integrated tests of the CSM and LV arebeing performed, checks will be made on the Lab electronic subsystems, and a pres-sure decay test made on the RCS. EMC checks will be performed on the Lab duringthe CSM/LV "plugs out" test. The Launch Escape System (LES) will be installed andML swing arm tests will be performed. The space vehicle will then be moved to thelaunch pad.

3.3-3.3-7 Launch Pad 39 (Saturn V). The space vehicle will be transported to thelaunch pad on the ML and placed in position. Range/Space vehicle RFI tests will beperformed before the Mobile Service Structure (MSS) is moved into position. Afterthe MSS is positioned a checkout of the Lab subsystems will be performed. Lab RFSubsystem checks will be performed using antenna hats and SLA reradiating antennas.Following subsystems verification, the Lab will participate in a flight readinesstest and then hypergolic servicing will be accomplished. A countdown demonstrationwill be performed to verify the capability of accomplishing terminal countdown pro-cedures in the allotted time. EPS flight batteries will be installed. After afinal subsystems checkout, all ACE carry-on equipment will be removed from the SLA.Servicing of the fuel cell reactant cryogenic storage section and start up of thefuel cells will be accomplished during the final countdown. Lab pyrotechnic initi-ators and pyrotechnic batteries will then be installed. Final fluid servicing willbe accomplished, and the Lab cabin will be closed out and pressurized. All GSEwork platforms will be removed from the SLA, and the SLA will be closed out.

3.3o3.3»8 Launch Pad 3TB (Saturn IB). The spacecraft will be stacked on the •Launch Vehicle (LV) at Launch Pad 37B. The Launch Escape System (LES) will beinstalled and launch tower swing arm tests will be performed. While integratedtests of the CSM and the LV are being performed, the Lab subsystems will bechecked. The service structure will be removed and Range/Spacecraft RFI testsand "plugs out" tests performed. The service structure will be repositioned andpressure decay tests made on the Lab RCS. Fuel cells will be started and opera-tionally checked using ambient reactants, - shut down (if required), and thenprepared for launch. A countdown demonstration will be performed to verifythe capability of accomplishing terminal countdown procedures in the allotted time.The Lab will then participate in the Space Vehicle Simulated Flight. All hypergolic

(XVl)3-34

servicing will then "be accomplished and EPS flight batteries will be installed.After a final subsystems checkout, all ACE carry-on equipment will be removedfrom the SLA. Final fluid servicing including fuel cell reactant loading will beaccomplished and the Lab cabin will be closed-out and pressurized. Pyrotechnicinitiators and batteries will then be installed. All GSE and work platforms willbe removed from the SLA and the SLA will be closed out.

3.3-4 Shelter (Fig. 3-3-1C)


3.3.U.1.1 Hypergolic Test Building (West Cell). Upon arrival at KSC, the ascentstage will be transported to the West Cell of the HTB where the receiving inspectionof the stage will be performed. Checkout and servicing of the ECS Heat TransportSection will be performed to provide coolant capability during subsequent tests.Leak and functional checks will be performed on the RCS, Oxygen Supply and CabinPressurization Section, and Water Management Section plumbing. The ascent stagewill then be transported to the MSOB.

3.3-4.1.2 Manned Spacecraft Operations Building. After the ascent stage is trans-ported bo the MSOB, it will be installed in the cleaning fixture, inverted and thenmoved to the docking fixture, for the Shelter-CSM docking test. Upon completion ofthe docking test, the ascent stage will be moved to the cleaning fixture for re-inversion and Water Management, Heat Transport, and Oxygen Supply and Cabin Pres-surization Sections checkout. The ascent stage will then be prepared for matingwith the descent stage.


3-3=4.2.1 Hypergolic Test Building (East Cell). The descent stage will be trans-ported to the East Cell of the HTB where a receiving inspection will be performed„Leak and functional checks will be performed on the propulsion sections. Thesupercritical helium section will be serviced, and a heat leak check performed toestablish insulation integrity. The descent stage will then be transported to theRFSTF.

3.3.4»2.2. RF Systems Test Facility. Upon arrival of the descent stage at theKFSTF, non-flight landing gear will be installed and the stage mounted on the three-axis positioner. The landing radar will be boresighted by comparing electrical beamposition with the optical normal. The descent stage will then be removed from thethree-axis positioner, the landing gear removed, and the stage transported to theMSOB.

3.3.4.2.3 Manned Spacecraft Operations Building. The descent stage will be trans-ported to the SLA stand for a descent stage/SLA fit check. The stage will then bemoved to the proper work stand for mating with the ascent stage. The Heat Trans-port and Water Management Sections of the ECS will be checked prior to mating thedescent and ascent stages.

3«,3o403 Mated Shelter

30-3.4.3.1 Manned Spacecraft Operations Building. The ascent stage and descentstage will be mated and permanent interstage plumbing connections leak checked.The Shelter will then be transported to the FCSTF.

(XVl)3-35

3.3. .3.2 Fuel Cell System Test Facility. The Shelter vill be installed on acheckout stand. Electrical functional checks vill be performed on the reactantstorage section, including the associated controls and displays. After leakageand mechanical functional checks are performed, the storage section will be ser-viced vith high-pressure gaseous reactants. Flow checks will be performed withoutstart-up of the fuel cells. The vehicle will then be deserviced and transported tothe MSOB.

3-3-^.3-3 Manned Spacecraft Operations Building. Upon arrival at the MSOB theShelter will be -installed on a checkout stand and the Heat Transport Section willbe leak checked and serviced. The Power Distribution Section and InstrumentationSubsystem capability to support vehicle testing will be verified. Checks will bemade to verify Optical Tracker calibration and Landing Radar (LR) antenna andS-band antenna alignments. Electrical and electronic subsystem tests and GWCSpolarity verification will be performed. The up-data link will be functionallychecked.

An integrated GWCS test will be performed, wherein all functional flow pathsand modes of operation are verified. To demonstrate the ability of the electricaland electronic subsystems to work together as a system, a mission oriented "plugsout" test will be performed. The Shelter will then be installed in the altitudechamber.

After installation in the altitude chamber, ECS, cabin, airlock, and hatch leakageand functional tests will be performed prior to evacuating the chamber. Thechamber will be evacuated with the Shelter unmanned, and ECS functional tests willbe performed to demonstrate the capability of the ECS to support manned altitudetests. The structural integrity of the Shelter, during emergency re-pressurizationof the chamber, will be demonstrated in the unmanned altitude run. The mannedaltitude tests will demonstrate ECS and Crew Provisions capabilities with a manin the loop. The fuel cells will be operated in the altitude chamber concurrentlywith ECS checkout. Electrical power will be generated to other subsystems,simulating critical power distribution requirements, and redundant and emergencymodes of electrical control logic. In addition, the RTG and associated heat pipeoperation will be functionally checked. After the manned altitude tests, thevehicle will be deserviced and transported to the RFSTF.

3-3-^-3-^ Radio Frequency Systems Test Facility. The vehicle will be mounted onthe RFSTF three-axis positioner. A communications test with the MSFW will beperformed to verify the compatibility between the Shelter Communications Sub-system and MSFN. The S-band steerable antenna will be functionally tested,utilizing an S-band source located at some distance from the vehicle and motioninputs to the positioner. Shelter EMC and Shelter/CSM EMC tests will be performedat the RFSTF. Upon completion of testing at the RFSTF, the vehicle will betransported to the MSOB.

3.3. .3.5 Manned Spacecraft Operations Building. The Shelter will be installedon a landing gear installation fixture in the MSOB. Flight landing gear will beinstalled and functionally checked. Pyrotechnic circuitry will be verified andthe explosive devices (minus initiators) mechanically installed. After thermalshield installation, radiator servicing, and exterior refurbishment, the Shelterwill be mated with the GSM/SLA in the SLA stand. The spacecraft (Shelter/CSM/SLA)will be transported to the VAB after SLA work platform installation and top deckbuildup.

(XVI)3-36

3.3.^-3-6 Vehicle Assembly Building. The Spacecraft (S/C) will be stacked onthe Launch Vehicle (LV) which will be in place on the Mobile Launcher (ML) in ahigh-bay area of the VAB. While integrated tests of the CSM and LV are beingperformed, checks will be made on the Shelter electronic subsystems, and pressuredecay tests made on the RCS and Descent Propulsion Subsystem. EMC checks will beperformed on the Shelter during the CSM/LV "plugs out" test. The Launch EscapeSystem (LES) will be installed and ML swing arm tests will be performed. Thespace vehicle will then be moved to the launch pad.

3-3-^-3..7 Launch Pad 39 (Saturn V). The space vehicle will be transported tothe launch pad on the ML and placed in position. Range/Space Vehicle RFI testswill be performed before the Mobile Service Structure (MSS) is moved into position.After the MSS is positioned, a checkout of the Shelter subsystems will be performed.Shelter RF subsystem checks will be performed, using antenna hats and SLA re-radiating antennas. Fuel cells will be started and operationally checked usingambient reactants, shut down, and then prepared for launch. Following subsystemsverification, the Shelter will participate in a flight readiness test and thenhypergolic servicing will be accomplished. A countdown demonstration will beperformed to verify the capability of accomplishing terminal countdown proceduresin the allotted time. EPS flight batteries will be installed. After a finalsubsystems checkout, all ACE carry-on equipment will be removed from the SLA.Shelter pyrotechnic initiators and pyrotechnic batteries will then be installed.Final fluid servicing including fuel cell reactants will be accomplished, and theShelter cabin will be closed out and pressurized. The RTG will be fueled, allGSE work platforms removed from the SLA, and the SLA will be closed out.

3-3-5 Taxi (Fig. 3-3-1D)


3.3.5.1.1. Hypergolic Test Building (West Cell). Upon arrival at KSC the ascentstage will be transported to the West Cell of the HTB where -the receiving in-spection of the stage will be performed. Checkout and servicing of the ECS HeatTransport Section will be performed to provide coolant capability during sub-sequent tests. Leak and functional checks will be performed on the RCS, AscentPropulsion, Oxygen Supply and Cabin Pressurization Section, and Water ManagementSection plumbing. The ascent stage will then be transported to the MSOB.

3.3.5.1.2 Manned Spacecraft Operations Building. After the ascent stage istransported to the MSOB it will be installed in the cleaning fixture, invertedand then moved to the docking fixture, for the Taxi-CSM docking test. Uponcompletion of the docking test, the ascent stage will be moved to the cleaningfixture for inversion and Water Management, Heat Transport, and Oxygen Supplyand Cabin Pressurization Sections checkout. The ascent stage will then be pre-pared for mating with the descent stage.


3.3.5.2.1 Hypergolic Test Building (East Cell). The descent stage will betransported to the East Cell of the HTB where a receiving inspection will be per-formed. Leak and functional checks will be performed on the Propulsion Sections.The supercritical helium section will be serviced, and a heat leak check per-formed to establish insulation integrity. The descent stage will then betransported to the RFSTF.

(XVl)3-37

3.3.5.2.2 RF Systems Test Facility. Upon arrival of the descent stage at theRFSTF, non-flight landing gear will be installed and the stage mounted on athree-axis positioner. The landing radar will be boresighted by comparing elec-trical beam position with the optical normal. The descent stage will then beremoved from the three-axis positioner, the landing gear removed, and the stagetransported to the MSOB.

3.3.5.2.3 Manned Spacecraft Operations Building. The descent stage will betransported to the SLA stand for a descent stage/SLA fit check. The stage willthen be moved to the proper work stand for mating with the ascent stage. TheHeat Transport and Water Management Sections of the ECS will be checked prior tomating the descent and ascent stages.

3.3.5.3 Mated Taxi

3.3.5.3.1 Manned Spacecraft Operations Building. After mating, the Heat TransportSection will be leak checked and serviced. The Power Distribution Section and In-strumentation Subsystem capability to support vehicle testing will be verified.Checks will be made to verify Alignment Optical Telescope calibration and Rendez-vous Radar (RR) antenna, Landing Radar (LR) antenna, and S-band antenna alignments.Electrical and electronic subsystem tests and GWCS polarity verification will beperformed.

An integrated GNCS test will be performed wherein all the functional flow pathsand'modes of operation are verified. To demonstrate the ability of the electricaland electronic subsystems to work together as a system, a mission oriented "plugsout" test will be performed. The Taxi will then be installed in the altitudechamber.

After installation in the altitude chamber, ECS, cabin, and hatch functional andleakage tests will be performed prior to evacuating the chamber. The chamber willbe evacuated with the Taxi unmanned, and ECS functional tests will be performed todemonstrate the capability of the ECS to support manned altitude tests. The struc-tural integrity of the Taxi, during emergency re-pressurization of the chamber,will be demonstrated in the unmanned altitude run. The manned altitude tests willdemonstrate ECS and Crew Provisions capabilities with a man in the loop. Inaddition, the RTG and associated heat pipe operation will be functionally checked.After the manned altitude tests, the vehicle will be deserviced and transported tothe RFSTF.

3.3.5.3.2 RF Systems Test Facility. The vehicle will be mounted on the RFSTFthree-axis positioner. A communications test with the MSFN will be performed tover.ify the compatibility between the Taxi Communications Subsystem and MSFN. TheS-band steerable antenna will be functionally tested, utilizing an S-band sourcelocated at some distance from the vehicle and motion inputs to the positioner.An ACS dynamic test will be performed with the use of the three-axis positioner.The vehicle will be oriented to various positions so that a gross determinationof ASA scale factors can be made. The Taxi will be optically aligned to the signalsource at the far end of the range and Rendezvous Radar pointing accuracy andtracking tests performed. At the completion of testing at the RFSTF, the vehiclewill be transported to the MSOB.

j/uvnmarL

(XVl)3-3&

3.3-5-3-3 Manned Spacecraft Operations Building. The Taxi will be installed ona landing gear installation fixture in the MSOB. Flight landing gear will beinstalled and functionally checked. Pyrotechnic circuitry will be verified andthe explosive devices (minus initiators) mechanically installed. After thermalshield installation and exterior refurbishment, the Taxi will be mated with-theCSM/SLA in the SLA stand. The spacecraft (Taxi/CSM/SLA) will be transported tothe VAB after SLA .work platform installation and top deck buildup.

3.3.5.3.^ Vehicle Assembly Building. The Spacecraft (S/0) will be stacked onthe Launch Vehicle (LV) which will be in place on the Mobile Launcher (ML) in ahigh-bay area of the VAB. While integrated tests of the CSM and LV are being per-formed, checks will be made on the Taxi electronic subsystems, and pressure decaytests made on the RCS and Propulsion subsystems. EMC checks will be performed onthe Taxi during the CSM/LV "plugs out" test. The Launch Escape System (LES) willbe installed and ML swing arm tests will be performed. The space vehicle willthen be moved to the launch pad.

3.3-5-3-5 Launch Fad 39 (Saturn V). The space vehicle will be transported tothe launch pad on the ML "and placed in position. Range/Space Vehicle RFI testswill be performed before the Mobile Service Structure (MSS) is moved into position.After the MSS is positioned, a checkout of the Taxi subsystems will be performed.Taxi RF subsystem checks will be performed using antenna hats and SLA reradiatingantennas. Following subsystems verification, the Taxi will participate in a flightreadiness test and then hypergolic servicing will be accomplished. A countdowndemonstration will be performed to verify the capability of accomplishing terminalcountdown procedures in the allotted time. EPS flight batteries will be installed.After a final subsystems checkout, all ACE carry-on equipment will be removed fromthe SLA. Taxi pyrotechnic initiators and pyrotechnic batteries will then be in-stalled. Final fluid servicing will then be accomplished, and the Taxi cabin willbe closed out and pressurized. All GSE work platforms will be removed from theSLA, and the SLA closed out.

Working Days(2-Shift, 5-Day Week)

A

I

Phase I LabSaturn IB

Phase I LabSaturn V

Phase II LabSaturn IB

rB

Phase II LabSaturn V

ShelterSaturn V

D TaxiSaturn V

Notes: 1. Cross-hatched portions representperiods of ACE utilization.

2. For each cross-hatched area, the "%"figure represents the estimated portionof the two-shift "checkout" time whichwill require ACE.

RFSTF Desc Only, LR Boresignt

48 44

60%

40 36

60%'RFSTH-MSOB-

32 28 24 20 16 12

Launch Pad 37B100%

60%•RFSTF- MSOB

50%VAB

100%• Launch Pad 39 •

Lab TopCSM I DeckSLA I Build-Mate | up

I

Launch Pad 37B100%

Co mmTests1

LabSLA

Mate

75%MSOB

60%-RFS -MSOB-

50%VAB

100%- Launch Pad 39 •

Lab I TopCSM I DeckSLA Build-Mate I

Iup

(XVI) 3-39^0

Fig. 3.3-1 KSC Facility/AES VehicleFlows

CONFIDENTIAL

3-4 FRELAUWCH OPERATIONS SCHEDULE

The integrated -schedule of AES-LEM prelaunch operations at KSC is presented inFi-g. 3.4-1. The schedule is made up of the operations flows of each of the fourbasic AES-LEM vehicles shorn in Fig. 3-3-1 and discussed in Paragraphs 3.3-2(Phase I Lab), 3.3-3 (Phase II Lab), 3-3-4 (Shelter), and 3-3-5 (Taxi). The over-all flow is based on the launch schedule of Ref. 4-7 and maximum use of existing andcurrently planned Apollo facilities and GSE. The KSC modification time for thePhase I Labs is shown for reference only. Discussion of manufacturing and modifi-cation time is found in the Manufacturing Plan, Vol XVII. The Apollo LEM checkoutspans are shown for reference purpose only, and are based on Ref. 4-6.

3-4.1 Assumptions

The preliminary prelaunch schedule and planning analysis reflects a number of basicassumptions. The major assumptions affecting vehicle and checkout facilityscheduling are as follows:

a. For planning purposes the Launch Complex (LC) assignments are:Saturn V Flights: LC 39Saturn IB LEM Lab Flights: LC 37BSaturn IB CSM Flights: -LC 34

b. The checkout times shown for LC 37B and 39 are based on Launch Vehicle (LV)checkout requirements. The spacecraft checkout (CSM/AES LEM) is performedconcurrently with LV operations.

c. Checkout scheduling is based on a two-shift work-day, 5-d.ay work week, and22-work-day month.

d. To avoid simultaneous checkout operations (lOO$> overlap) between two AES-LEM spacecraft, and to avoid simultaneous launches of two manned spacecraft,the dual launch rendezvous missions are scheduled so that the Lab islaunched at the end of the month and the respective CSM is launched at thebeginning of the next month as follows:

214 LEM Lab I End of June 1969215 CSM Beginning of July 1969

216 LEM Lab I End of October 1969217 CSM Beginning of November 1969

218 LEM Lab II End of April 1970219 CSM Beginning of May 1970

220 LEM Lab II End of June 1970221 CSM Beginning of July 1970

224 LEM LAB II End of February 1971225 CSM Beginning of March 1971

CONFIDENTIAL &uwnan.

(XVI)3-U2 CONFIDENTIAL

For the same reasons, the triple launch mission is scheduled as follows:

523 CSM/LEM Lab II Mid August 1971228 LEM Lab II End of September 1971229 CSM Beginning of October 1971

Saturn V launches (with the exception of 523) are assumed at the beginningof the month.

e. As a goal, all vehicle/experiment comprehensive testing will be accomplishedin the Manned Spaceflight Operations Building (MSOB). However, it isrecognized that specific exceptions may require other facilities such asthe Hypergolic Test Building (HTB), and the RF Systems Test Facility(EFSTF).

f. All experiments are installed before integrated testing.

g-. It is assumed that the Phase I Labs-have not been cycled through any partof the prelaunch checkout flow before the start of KSC modification.

h. The requirement for modification of the Phase I Labs at KSC necessitatesthe demonstration of vehicle subsystem compatibility while operating as asystem. This will be accomplished in the MSOB in three parts:

Integrated Subsystem VerificationElectro-Magnetic Interference (EMl) VerificationMission Oriented Tests

The total Phase I Lab MSOB checkout will be longer than the Phase II LabMSOB checkout as a result of this integrated subsystems compatibilityverification. (For the Phase II Lab, this will have been accomplished be-fore the Lab leaves Bethpage).

i. Stage mating can only be performed in the MSOB.

j. Weight and eg of the Phase I Labs will be determined during the manufac-turing modification period at KSC. Weight and eg of the Phase II Labs willbe determined at Bethpage.

CONFIDENTIAL

1967 1968

J F M A M J J A S O N D

Veh., Fit., LaunchComplex

LEM-3,'503,39

LEM-4, 504,39

LEM-5, 505I I

LEM-6,506,L.O. Lab,

LEM & PHASE I LABS

PHASE H LABSSHELTERS & TAXIS

Prelaunch Checkout Time Spans

AESS/C

.Phase ILab

Phase IILab

Shelter

Taxi

SaturnLaunchVehicle

IBV

IBV

WorkJDays

8378

9388

87

81

CONFIDENTIAL (XVl)3-U3/iU

J A S O N DJ F M A M JJ A S O N DJ F M A M J J A S O N D J F M A M J

Facility CodeHypergolics TestBuilding

Manned Spacecraft I IOperations Building

RF Systems TestFacility

Fuel Cell SystemsTest Facility

Vehicle AssemblyBuilding

ISM ,215,34

StorageAreaLEM-9, 512,

ModificationPeriodCSM ,217,34

E.G. Lab, 513,39

Shelter, 514, 39i I

Taxi, 515,

E.O. Lab, 218,i i i i iCSM ,219,34

I I i i

a) Launch dates based on ML 65-1b) CSM & basic LEM time spans

are shown for reference onlyc) Bar charts are drawn using 22

working days in a monthd) No interface with CSM flows

is indicatede) The vehicle is always in a demoted

configuration while in the HTB

E.O. Lab, 516,39

E.b.Lab,'22C),37Bi i i i

CSM ,221,34I I i I

L. O. Lab, 517,39

E.O. Lab, 518

Shelter, 519,

Taxi, 520,

E.O. Lab, 224, 37BI I I I ICSM ,225,34

I I I IE.O. Lab, 521,39

i i iL. O. Lab, 522,

E.O. Lab, 523,I I I I IE.O. Lab, 228.37B

CSM ,229,34i i i

Shelter, 524,39i i

Taxi. 525. 39

Fig. 3.4-1 Prelaunch Operations Schedule

CONFIDENTIAL

3-5 CHECKOUT REQUIREMENTS AT KSC

The checkout operations for the AES-LEM vehicles (Paragraph 3-3) will use the exist-ing Saturn/Apollo checkout facilities, Acceptance Checkout Equipment (ACE) andGround Support Equipment (GSE). Additional requirements for checkout are the re-sult of the high launch rates (Paragraph 3«*0 and "the AES-LEM modifications to thebasic LEM configuration.

This paragraph describes the existing checkout facilities and discusses the facil-ity, ACE, and GSE requirements resulting from the AES-LEM configurations and launchschedule.

3-5-1 Facility'Description

The AES-LEM checkout functions are performed in the facilities of the KSC Industri-al area and Saturn Launch complexes, specifically:

a. Manned Spacecraft Operations Building (MSOB)b. Hypergolics Test Building (HTB)c. RF Systems Test Facility (RFSTF)d. Fuel Cell Systems Test Facility (FCSTF)e. Launch Complex 39 (Vehicle Assembly Building and Launch Pads) or Launch

Complex 37 (Launch Pad 37B).

3.5.1.1 Manned Spacecraft Operations Building (MSOB)

The MSOB, with the LEM addition, will be the central area for Apollo LEM checkoutand maintenance, and will provide space for engineering and administration person-nel. The major areas within the building are:

a. ACE and Data Facility - Data Processing (off-line) and ACE Stations(Computer Rooms, Control Rooms, and Terminal Facility Rooms)

b. Altitude Chamber - Simulates altitude conditions for spacecraft systemscheckout. The chamber is capable of achieving an altitude of 250,000 ftwithin 1 hour, returning to sea level in 150 sec. in an emergency, andreturning to sea level from 250,000 ft in 16 min-normal operation.

c. Assembly and Test (A&T) Area - Polarity fixture, A/S & D/S work stands,and checkout equipment for LEM system checkout

d. Instrumentation Calibration Laboratory - Calibration of flight and testinstrumentation

e. Integrated-Test Area (ITA) - (SLA stand and landing gear installation fix-ture) fit checks, stacking, and landing gear installation

f. Maintenance Areas - Laboratories and bench maintenance equipment for sched-uled and unscheduled maintenance:

- Battery LaboratoryCommunications/Radar Laboratory

- ECS Laboratory- EPS Laboratory

• - Modification Shop- S&C Laboratory

g. Photography Laboratory - Photographic processing supporth. Propellant Gaging Laboratory - Calibration & maintenance of propellant

quantity gaging systems (PQGS)

(XVI) 3-*t 6

i. S-Band and UHF Electronics Laboratory - Support of scheduled open-loop(RF link) and closed-loop (hardline) checkout tests of LEM communications.

j. Spare Part and Tool Room - Hand tools required for LEM checkout andshelf-type spare parts.

3.5.1.2 Hypergolics Test Building (HTB)

The facility has two test cells (East Cell and West Cell) for conducting LEM AscentPropulsion, Descent Propulsion, and RCS functional and leakage checks, and gascold-flow operation. The facility has the capability for supporting substitute-propellant cold-flow operations. The control room in the HTB affords a centralizedcontrol station for directing tests in the East and West Test Cells, provides con-tinuous data recording equipment and ACE interleaving equipment for Propulsion,RCS, and supporting subsystems under test.

3.5.1.3 RF Systems Test Facility (RFSTF)

The RFSTF will provide the facilities required to perform test and alignment checksof AES-LEM radiative equipment. The facility consists of a two-story controlbuilding and a 50-ft wooden tower, 1000 ft away from the control building, with aspecially prepared strip between the two. The Control Building houses most of theGSE required to perform the RF subsystem tests. A three-axis positioner for theLEM vehicle will permit orientation of the AES-LEM antennae .toward the tower at theopposite end of the range, while a control console in the building will permit re-mote control of antenna position at the down-range tower.

The strip between the Control Building and tower is a wide grass strip, gradedlevel to control ground reflections. The strip and surrounding area will be freeof any signal reflecting discontinuities.

The Range Tower is a rigid, wooden structure, used to support receiving and trans-mitting antennae. Optical targets used for sighting and calibration of AES-LEMantennae and GNCS components will also, be located here. Signal sources for theradars will be housed at the tower and feed horns will be mounted on a movablecarriage, the position of which is servo controlled from the Control Building.

3.5.1A Fuel Cell Systems Test Facility

This facility is not used in the basic LEM prelaunch checkout. However, for thePhase II Lab and Shelter configurations using fuel cells for electrical power gen-eration, an existing FCSTF will provide the area for checkout of the fuel cell andreactant storage systems.

An L0p pad and LHp pad for reactant servicing equipment are located at a safe dis-tance from the test cell and from each other. The facility incorporates a fuel-cell pre-installation acceptance laboratory to provide the capability for fuel cellreplacement and unscheduled maintenance. An equipment room will be provided forprotection of GSE and ACE items which are not designed to 'operate in a hydrogen-hazardous environment.

3.5.1.5 Launch Complex 39 (VAB and Pad)

3.5.1.5.! Vehicle Assembly Building (VAB). The VAB will permit simultaneous stack-ing of four Apollo spacecraft to Saturn V launch vehicles in a vertical position.Buildings within the main structure will provide individual office and work area

(XVI)3- 7

for each space vehicle assembly. Enclosed, air conditioned, work platforms thatmove vertically along the inside walls of the VAB telescope in and out to permitworking on any point of the space vehicle. Heavy duty cranes and hoists within theVAB will facilitate stacking operations. The fully assembled Space Vehicle (S/V)and Mobile Launcher (ML) will be moved intact in a vertical position to the launchpad on a tractor-crawler.

3.5.1.5.2 Launch Pads 39A and 39B. The S/V will be transported to the Launch Padon the ML and placed in position on the pedestal. A work platform will be providedon the Mobile Service Structure (MSS) at the LEM level. This work platform will beused to gain access through the SLA to the LEM, and it will also be used to supportGSE that must be near the spacecraft. Prior to launch, the MSS is moved away fromthe space vehicle and ML by the tractor-crawler. The ML is designed to withstandthe launch environment.

3.5.1.6 Launch Complex 37B

As part of the basic Apollo Program, Complex 37B will be configured to support aSaturn IB launch of a LEM and CSM. Pad 37B is equipped with a pedestal, MobileService Structure umbilical tower and ground control station. This facility willbe utilized for the Phase I and II Lab-CSM Boilerplate flights. (Pad 3^, as in theApollo Program, will be in the configuration to support the manned CSM launcheswhich rendezvous with the unmanned Saturn IB-launched Labs.)

3.5.2 Facility Requirements

The preliminary facility requirements are based on the prelaunch checkout opera-tions discussed in Paragraph 3-3 and the schedule presented in Paragraph 3.U. Therequirements for the facilities described in Paragraph 3.5.1 are discussed in thefollowing paragraphs.

3.5-2.1 Assumptions

a. The overlap in KSC prelaunch checkout operations for the different AES-LEMvehicles is based upon the following nominal checkout time spans (Paragraph3.3) and KSC modification time spans (Vol. XVII):

Time (months)- Basic LEM C/0 3-2/3- Phase I Lab Modification at KSC- Phase I Lab C/0 (Sat. IB) 3-- Phase I Lab C/0 (Sat. V) 3-1/2- Phase II Lab C/0 (Sat. IB) k-l/k- Phase II Lab C/0 (Sat. V) h- Shelter C/0 k- Taxi C/0 3-2/3

b. Storage facility is not identified as a particular building.c. The requirements for the Phase I Lab modification are assumed to be satis-fied :

- At a facility other than one which is presently planned for checkout, or- By modification (or expansion) of one of the existing or currently

planned KSC facilities.

(XVI)3-48

d. As part of the basic Apollo Program, Launch Pad 37E will be configured tosupport a Saturn IB launch of a LEM/CSM Boilerplate (Ref. 4-1.) and LaunchPad 34 will be configured to support a Saturn IB launch of a manned GSMwithout LEM.

e. The facility "turnaround" time relative to spacecraft operations is esti-mated to be 1 month for Launch Pad 37B and 15 days for Launch Pad 39.

f. All LEM and AES-LEM vehicles will be checked for EMC in the Bethpageshielded chamber before shipment to KSC.

g. Background EMI noise levels in the unshielded KSC facilities will be suffi-ciently high and variable to mask potential EMC problems.

3.5.2.2 General Requirements

LEM/AES-LEM vehicle loading through the KSC facilities is shown in Fig. 3.5-1.When only one stage (ascent or descent) is in a given facility, the figure indi-cates a half vehicle. The KSC facilities presently planned for the Apollo Program(Paragraph 3.5-1) will be used to satisfy the requirements of the AES-LEM checkoutwith the additions noted below.

3-5-2.2.1 Hypergolic Test Building (HTB). The HTB presently configured to supportan Apollo LEM ascent stage in one cell and a descent stage in the other cell willmeet the requirements of AES-LEM vehicle Propulsion and RCS checkout. The HTBloading shown in Fig. 3-5-1 indicates overlap of the Taxi ascent and descent stagesand Phase II Lab descent stage in December 1969 and October 1970. Since the overlapis approximately 1 week, the problem may be amenable to schedule adjustments.

3.5.2.2.2 RF Systems Test Facility (RFSTFj. The requirements for a RFSTF at KSC,as described in Paragraph 3-5.1 are based on the prelaunch checkout operations dis-cussed in Paragraph 3-2. The checkout operations that must be performed at such afacility because of subsystem requirements are:

a. RR pointing accuracyb. RR tracking accuracy and antenna stabilization-c. LR boresightingd. AGS dynamic testinge. S-band antenna functional testingf. X-Y scanner functional checkoutg. AES-LEM/CSM EMC checkouth. Required experiment checkout

3.5.2.2.3 Fuel Cell Systems Test Facility (FCSTF). This existing facility is notpresently required by the Apollo Program LEM. The Phase II Labs and Shelters willrequire the facility as described in Paragraph 3-5-1 for checkout of the fuel cellsand reactant storage systems. The checkout operations, as discussed in Paragraph3-3) require that the facility be capable of supporting:

a. Leakage and functional checks of fuel cell and reactant storage systems.b. Liquid hydrogen and liquid oxygen servicing of Phase II Lab reactant

storage.c. High-pressure gaseous hydrogen and oxygen servicing of Shelter reactant

storage.d. Functional flow checks of the fuel cells systems.

(XVI)3-U9

As presently planned, the fuel cells will not be started in the FCSTF as part of thefuel cell checkout operation in this facility. The fuel cells will be operated inthe Altitude Chamber (MSOB) and Launch Pad.

3.5.2.2.U Manned Spacecraft Operations Building (MSOB). The MSOB, as described inParagraph 3-5-1^ requires an addition of an EMC shielded chamber and modification ofthe altitude chamber. The shielded chamber is necessary to demonstrate the EMCcharacteristics of each Phase I Lab after KSC modification and integration of ex-periments. This recommendation is based on:

a. The relatively high and variable ambient EMI levels at the KSC testingfacilities.

b. Low Lab communication circuit margins under operating conditions.c. Detection and correction of EMC anomalies not determined in the development

program due to manufacturing differences.d. Verification of the absence of all EMC anomalies that were detected during

development testing.

Ambient EMC testing is unacceptable due to the high and variable EMI levels thatwill be found at the KSC test facilities. This background EMI will mask some of theEMC problems while inducing others. The result of EMC testing in an uncontrolledenvironment will be an increase in testing time and the consequent possibility ofschedule impact without any further assurance of having a compatible vehicle. ThePhase I Lab EMC testing is in line with the present LEM test approach and the AEStest philosophy for Phase II Labs, Taxi, and Shelter since these three vehicles willundergo EMC testing in a shielded chamber at Bethpage.

The altitude chamber requires modification so that the fuel cells of the Phase IILab and Shelter can be started and operated using facility supplied gaseousreactants.

Fig- 3«5~1 indicates that, for short periods of time (l week or less), three vehi-cles are in the MSOB at one time; and that, for periods of 1 to 2 weeks, 'two vehi-cles are in the MSOB at one time. In all of these instances, no two vehicles re-quire the same checkout area at the same time except during AES-LEM/SLA fit checkand AES-LEM/CSM mating functions which are normally accomplished in the SLA standlocated in the GSM checkout area. This potential problem may be solved by theeventual elimination of the requirement for the fit check or by relocating equip-ment, necessary for the fit check, from the SLA stand to another available stand.

3.5.2.2.5 Launch Complex. Since no more than one AES-LEM vehicle is in the VAB atone time, no changes to the facility, as presently planned for the Apollo Program,are required. The time span between LC 39 single-pad operations is greater than 30days with two exceptions: between the launch of 52^ Shelter and pad arrival of 515Taxi, and between the launch of 52^ Shelter and pad arrival of 525 Taxi. The timespans involved are less than 15 days; however, the two Saturn V launch pads at Com-plex 39 can satisfy the launch schedule. The time span between the launch of 218Lab and arrival at the pad of 220 Lab is less than 1 month and, therefore, an areato be further studied. Since significant overlap in operations exists between thelaunch pad of Complex 37 and the launch pad( s) of Complex 39> sharing of G-SE betweenthese complexes will be limited.

CXVI)3-50

3-5.3 Acceptance Checkout Equipment (ACE) Requirements

A requirement- for four ACE stations to support the AES-LEM vehicles is indicated bythe density of vehicles at KSC; i.e., by combination of pre-launch checkout time,launch schedule, and different vehicle/mission combinations of the Apollo LEM,Phase I Lab, Phase II Lab, Shelter, and Taxi.

3-5.3-1 Assumptions

a. No variations in the planned launch rates nor vehicle flows are considered,"b. Wo contingencies are considered.c. Validation of KSC ACE programming is performed at the KSC ACE facilities.d. A nominal ACE turnaround time of 3 weeks is assumed for station reconfigu-

ration and internal checkout, software verification, and checkout of ACEperipheral equipment.

e. One ACE station is capable of supporting only one vehicle at one facility(or test station) at one time.

f. When a facility-vehicle requirement is within the program and configurationof an ACE station, switching time between facilities or between vehicles isnegligible.

g. While an ACE station is supporting launch pad operation, it will not beused for any other operations.

h. Wo consideration is given to concurrent sharing of an ACE station between"LEM" and "CSM" function.

i. Each vehicle will have its own set of ACE carry-on equipment.

3.5-3.2 Requirements

Two different viewpoints were considered in the determination of requirements:

a. Orienting the ACE station(s) to satisfy the general requirements estab-lished by facility ( or group of facilities) usage and,•

b. Orienting the ACE station(s) to satisfy the general requirements estab-lished by a vehicle/mission combination.

The ACE loading presented, one of many possible combinations of a and b, illus-trates the nature of the problems that may arise in optimizing ACE utilization.The indication, nevertheless, of a total ci four ACE stations for Grumman use, isapparent throughout the loading study.

Fig. 3-5-2 presents an estimated percentage of ACE utilization in each facilityduring the KSC checkout cycle for each of the AES vehicles. The estimates werebased on the present LEM Ground Operations Requirements Plan (Ref. U-6) and thevehicle flows presented in Paragraph 3-3. The cross-hatched portions of each flowin Fig. 3-5-2 represent periods of ACE utilization, whereas the "open" areas indi-cate no ACE utilization. The "percentage figures" for each cross-hatched arearepresent the estimated portion of the two-shift day "checkout" time which will re-quire ACE. This figure does not assume ACE setup time nor does it assume the dis-tribution of ACE utilization within the open area. Third-shift operation isreserved for ACE maintenance.

Fig. 3.5-3 presents the ACE station utilization in relation to the KSC prelaunchoperations and manufacturing schedule of the LEM and AES-LEM vehicles. The ACE

(XVI)3-51

utilization (approximately 1 month) during the KSC manufacturing cycle of Phase ILabs, based on the Manufacturing Plan, is assumed for subsystem build-up verifica-tion. Each ACE station is identified by a letter, and the utilization time isrepresented by the length of the arrows. The extension of the arrow to the left ofthe checkout cycle indicates the 3-week period for station "turnaround". Severaltimes in the flow, ACE sharing between vehicle/facility combinations is assumed(and noted in the figure) rather than requiring an additional ACE station.

Fig. 3,5-U-summarizes the ACE loading by relating it to the vehicle and facilityloading discussed in Paragraph 3.5-2. Utilization of four ACE stations is indi-cated from early 1969 through early 1971- Continuous four-station utilization isindicated from January to June of 1970. This ACE station loading is generated pre-dominantly by the density of vehicles and the different configurations of thesevehicles. The modification cycle for Phase I Labs requires that the additional ACEstation(s) are necessary sooner in the AES schedule time span than would be neces-sary if no modification at KSC were accomplished.

3-5-^ Ground Support Equipment (GSE) Requirements

The GSE requirements to implement the prelaunch checkout are based on the checkoutfunctions, AES vehicle configurations, and KSC facilities.

3.5. .1 Assumptions

a. The following activities are prerequisite to any given checkout function:

- Power distribution verificationInstrumentation turn on and verificationCabin entry (and closeout)Heat transport (HTS) servicing

b. One set of ACE carry-on equipment is with each vehicle'throughout thecheckout cycle.

c. Power 'supplies (ac and dc) are a function of facility.d. The pressure maintenance unit (all- AES-LEM vehicles), IMU portable tempera-

ture controller (Shelter and Taxi), and ASA portable temperature controller(Phase I Lab, Phase II Lab, and Taxi) are always with a given vehicle.

e. The Instrumentation Subsystem of the Phase I Labs will be calibrated in theLEM instrumentation lab area located in the MSOB.

f. Requirements generated by unscheduled maintenance, Phase I Lab manufacturingat KSC, and ACE utilization are not shown in Fig. 3«5-5«

g. The figure also does not indicate the order or sequence of checkout func-tions within a given facility nor the order in which the facilities will beused.

3.5-^.2 Equipment Requirements

Fig. 3-5-5 summarizes the basic requirement for ground support equipment generatedby the prelaunch operations at KSC for AES-LEM vehicles. These requirements, re-lated to each facility and to the checkout functions within these facilities, arepresented in a matrix format. Each major block, which is further subdivided intoquadrants, relates the GSE requirements to the vehicle, facility, checkout function,

(XVl)3-52

and estimated, extent of change from presently planned LEM GSE. The quadrantsrelate to the AES vehicles as follows:

a. Upper left-hand corner: Phase I Labb. Upper right-hand corner: Phase II Labc. Lower left-hand corner: Shelterd. Lower right-hand corner: Taxi

The "extent of change" of the required GSE compared to the presently planned LEMequipment is identified as follows:

m The checkout function is the same as presently planned for the current LEM,and the equipment which is planned to satisfy the LEM requirement willsatisfy the AES-LEM requirement.

[J The checkout function is basically the same as presently planned for thecurrent LEM; however, the equipment which is planned to satisfy the LEM re-quirements may have to be modified to extend its capacity; e.g., greatercapacity in fluid storage carts, greater load capacity for handling equip-ment and support stands, and/or greater heat sink capacity in cooling equipequipment.

[[ The checkout function is unique to AES-LEM configurations, i.e., to subsys-tems which have major additions or modifications.

3-5-^.3 Scheduling Requirements

Review of Fig. 3-5-5 indicates that the majority of the AES-LEM requirements may bemet by using the first two categories. New equipment (based on what is now avail-able for LEM) is required for checkout and servicing of: fuel cell systems (Mid1969), cryogenic systems (Mid 1969), RTG systems (Late 1969), CSM crew provisions(Late 1969). Radiator systems (Late 1969), beacon transponder (Early 1970).

The requirement for a shielded chamber at KSC for Phase I Lab EMC verification(Paragraph 3.5.2) necessitates the use of additional antenna for coupling RFenergy into the chamber. Standard EMI test equipment (i.e., field strength meters,etc.) will also be required for the proper conduct of the EMC verification testing.

The quantity of equipment required will be dependent on the functional requirmentsas indicated in the chart and the vehicle/facility density (Fig.3«5«l)- Two areasof particular concern, indicated by Paragraph 3-5-2, are the MSOB and Launch Com-plexes. The density of vehicles through these areas and turn-around time of thelaunch pads will limit the degree of sharing equipment between facilities, betweenvehicles and between launch complexes. The Mobile Launcher (ML) may be particu-larly affected in that another ML is required to be configured with duplicate setsof operational equipment to support the launches of the Shelter and Taxi missionsscheduled for the early part of 1970 and latter part of 1970 and latter part of1971 (Fig. 3-^-1).

Prelaunch operations analysis indicates that the following equipment is required inan operationally ready condition during the noted time periods:

a. Communications Checkout Stations to support three vehicles at one time (Mid

1969)•b. Radar checkout equipment to support Shelter-Taxi Scheduling (Early 1970).c. Power Distribution Section checkout equipment to support three vehicles

(Early 1969).

(XVI) 3-53/5*+

d. Ascent stage handling or mated-vehicle handling equipment to support three .AES vehicles at one time (Mid 1969).

e. Descent stage handling and transportation equipment to handle two AESvehicles at one time (Early 1969).

f. Fuel cell system checkout and servicing equipment to support fuel cellcheckout (Mid 1969).

g. Propulsion/RCS servicing and checkout equipment to support propulsion/RCScheckout in both cells of the HTB and to support propulsion servicing onLC 37B and LC 39 at one time (Mid 1969).

h. Fuel cell checkout and servicing equipment to support countdown on LP 3TBconcurrent with checkout operations in the altitude chamber and FCSTF (Early1970).

io Vehicle cooling equipment to support the vehicles in the MSOB, in the HTB,in the FCSTF and on LP 39 concurrently (Early 1970).

j. ECS servicing and checkout equipment for LP 37B, LP 39 and altitude chamberat the same time (Late 1969).

k. Pressure maintenance units to support each vehicle at all facilities.1. The IMU and ASA Portable Temperature Controllers to maintain continuous

thermal environment for the IMU and ASA at all times,m. Calibration equipment for the Instrumentation Subsystem of the Phase I Labs

at KCS will (Mid 1968).

1967 1968

TOTALLEM/AES-LEMVEHICLES AT KSC

STORAGE

MODIFICATION& TIME PRIORTO PRELAUNCHCHECKOUT

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LAUNCHPAD 37B

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'/ut/n/nan.

Working Days(2-Shitt, 5-Day Week)

rA

I

Phase I LabSaturn IB

Phase I LabSaturn V

Phase IISaturn

rB

IPhase II

SaturnLabV

ShelterSaturn V

92 8.8 84 8.0 76

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Notes: 1. Cross-hatched portions representperiods of ACE utilization.,

2. For each cross-hatched area, the "%"figure represents the estimated portionof the two-shift "checkout" time whichwill require ACE.

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(XVI)3-57^5

Fig. 3.5-2 Vehicle/Facility ACEUtilization

'Aiunsnon.

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Note

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(£}^-^

(A)®

„ , . . SharesVehicle A(,E w.th

E. O. Lab 507 LEM 6

L.O. Lab 511 E.O. Lab 214

E.O. Lab 216 L.O. Lab 511

E.O. Lab 220 E.O. Lab 516

L.O. Lab 517 E.O. Lab 220

Taxi 525 Shelter 524

While

507LEM 6511214216511220516517220525524

is in

MfgHTB & MSOBMfgMfgMfgHTB & MSOBHTB, MSOB & FCSTFMSOB & FCSTFHTB & MSOBMSOB & RFSTFHTB & MSOB !MSOB & RFSTF

CONFIDENTIAL- (XVl)3-59/f

1970 1971

J F M A M J J A S O N D J F M A M J J A S O N D

Taxi, 515,391-I i I I I

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1 3

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i i i i iE.G. Lab, 516,39 |l

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L. O. Lab, 517,39

E.G. Lab, 518,39i i i

Shelter, 519,39

Taxi, 520, 39I I I I

E.G. Lab, 224, 37BI i i

Facility CodeHy.pergolics TestBuilding

Manned Spacecraft IOperations Building I

RF Systems Test [Facility i

Fuel Cell SystemsTest Facility

Vehicle AssemblyBuilding

LaunchPad

StorageArea

Mfg. Cycle(ACE Utilization)

Each ACE station isindicated bv a letter.

Utilization time is indicatedby length of arrows.

Extension of arrows to theleft of checkout cycleindicates stationturnaround time.

K$a " H vmiI II I A ll^»l

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225,34Y i i

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L.O. Lab, 522, 3911-i i i I

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Fig. 3.5-3 KSC ACE Utilization andPrelaunch Operations

TOTALLEM/AES-LEMVEHICLE &ACE STA.USE AT KSC

STORAGE

MODIFICATION& TIME PRIORTO PRELAUNCHCHECKOUT

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(XVl)3-6lf62*

F M . A M . J . J i A i ' S O N D J F M A M J J A S O N D J F M A M J J A S O . N D

Vehicle Loading

ACE Loading

A ' M J ' J A ' S ' O ' N D J F ' M A ' M J J A S ' O N ' DF ' M A M J ' J A ' S O N ' D

Fig. 3.5-4 Vehicle/Facility/ACELoading Summary

^\x^ Equipment•v. ; Requirements

, Checkout \.Function ^-~

HTB

FCSTF

MSOB

A/S or D/SInstl & RemovalECS C/0& ServicingD/S SHe Leak.Funct'l,Servicing & Heat Leak CksD/S Prop Leak& Funct'l CksA/S Prop Leak& Funct'l CksRCS Leak& Funct'l CksMated VehInstl & RemovalECS C/0& ServicingReactant Stor Leak.Funct'l,Servicing & Heat Leak CksFCA Leakage& Funct'l CksInstl & Removal ofS.tages_or_Mate<Ly.e!i ,A/S or D/S Mating

SLA Fit Ck, Stackingi Pyro InstlLdg Gear Instl& Funct'l CksDocking Tests

Alignment

PDS C/0

Instrumentation C/0

CommunicationsFunctional C/0Radar C/0

Beacon Xponder C/0

AGS Funct'l C/0 &Gyro Null Bias CompensPGNCS Subsys C/0

Mission - OrientedSystems C/0GNCS Polarity Cks

Integrated GNCS C/0

Integrated Ph 1 LabSubsystem C/0Ph 1 Lab EMC Cks

ECS & Crew Prov C/0& Svcg (Alt Chmbr)FCA Operational Cks(Alt Chmbr)RTG Funct'l Cks(Alt Chmbr)Radiator C/0& Servicing

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Fig. 3.5-5 AES-LEM Checkout EquipmentRequirements (Sheet 1 of 2)

^~\_^ Equipment^v. Requirements

Checkout"1-^Function ^v.

RFSTF

m

LAUNCH PADS^7 fat 39TPads37B&39

ASS Dynam Funct'l& Gyro Null BiasCommunications

L VerificationRendez RadarB'sight & TrackingLdg RadarB'sizht & C/0PGNCS Subsys C/0

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Fig. 3.5-5 AES-LEM Checkout EquipmentRequirements (Sheet 2 of 2)

,/uvnman-

(XVI)4-1

4. REFERENCES

4.1 Grumman Report LPL-611-3: "LEM-1 Flight Module Detailed Test Plan MissionSA-206 A )", 1 Sept 1965.

4.2 Grumman Spec, LSD- 380- 1?:"LEM-MSFN S-Band System Signal P & I Specification", 7 June 1965.

4.3 NASA PSRD, "Apollo/ Saturn V Program Support Requirements", Vol I, 1 April1965.

4.4 NASA-MSC-COC-1: "Apollo Checkout Criteria", 22 Oct 1964 (Updated - Selectionof Checkout and Flight Performance Measurements for the Apollo Program) .

4.5 Grumman Report LED-540-40: "Summary of Prelaunch Checkout and In-Flight Mon-itoring/Checkout Criteria", 26 Oct 1965.

4.6 Grumman Report LPL-610-3C: "Lunar Excursion Module Ground Operation Require-ments Plan", June 1965.

4.7 NASA Schedule, ML 65-1: "Flight Mission Assignment Plan for AES Planning",7 Aug 1965.

5. ABBREVIATIONS & SYMBOLS

(xvi)5-i

A

AC

ACA

ACE

ADV

A/E

AELD

AES-.LEM

AEA

• AGC

ACS

ALT

AOT

ARS

A/S

ASA

ATCA

ATM

ATT

AUTO

C

CCS

CES

CG

c/oC02

corn1

cosCSM

C & W

CWEA

CWG

Analog Display

Alternating Current

Attitude Controller Assembly

Acceptance Checkout Equipment

Advisory

Ascent Engine

Ascent Engine Latching Device

Apollo Extension System - Lunar Excursion Module

Abort Electronic Assembly

Automatic Gain Control

Abort Guidance Section

Altitude

Alignment Optical Telescope

Atmosphere Revitalization Section

Ascent Stage

Abort Sensor Assembly

Attitude and Translation Control Assembly

Angle Track Module

Attitude

Automatic

Caution Light

Command Control Section

Control Electronics Section

Center of Gravity

Checkout

Carbon Dioxide

Control

Cosine

Command Service Module

Caution and Warning

Caution & Warning Electronics Assembly

Constant Wear Garment

(XVI)5-2

D

DC

D/E

DECA

D/S

DSEA

EGA

ECS

EDS

EED

EMC

EMI

EMU

ENGR

EPS

ER

EXCT

F

FCA

FCSTF

G & N

GFE

GNCS

GOX

G He

G02/GH2

GSE

GSFC

H2

H20

HeHI

HTB

HTS

ID

IF

Digital Display

Direct Current

Descent Engine

Descent Engine- Control Assembly

Descent Stage

Data Storage Electronics Assembly

Electrical Control Assembly

Environmental Control Subsystem

Electro-Explosive Devices Subsystem

Electro-Explosive Devices

Electro Magnetic Compatibility

Electro Magnetic Interference

Extra-Vehicular Mobility Units

Engineer

Electrical Power Subsystem

Error

Excitation

Flag Indicator

Fuel Cell Assembly

Fuel Cell System Test Facility

Guidance & Navigation

Government Furnished Equipment

Guidance Navigation & Control Subsystem

Gaseous Oxygen

Gaseous Helium

Gaseous Oxygen/Gaseous Hydrogen

Ground Support Equipment

Goddard Space Flight Center

Hydrogen

Water

Helium

High

Hypergolic Test Building

Heat Transport Section

Identification

Intermediate Frequency

(xvi)5-3

IGIMU

INTCON

IR

ISOL

KSC

L

LES

LC

LG

LGC

LiOH

LR

LNDG

LN2

I02

LP

LV -

MCC

MG

MPID

MID

ML

MSC

MSFN

MSOB

MSS

N2WAV & GUID

°20/F

OG

OSCP

OPS

OTS

OUT

Inner Gimbal

Inertia! Measurement Unit

Interconnect

Infra red

Isolation

Kennedy Space Center

Light

Launch Escape System

Launch Complex

Landing Gear

LEM Guidance Computer

Lithium Hydroxide

Landing Radar

Landing

Liquid Nitrogen

Liquid Oxygen

Launch Pad

Launch Vehicle

Mission Control Center

Middle Gimbal

Manifold

Middle

Mobile Launcher

Manned Spacecraft Center

Manned Space Flight Network

Manned Spacecraft Operations Building

Mobile Service Structure

Nitrogen

Navigation & Guidance

Oxygen

Oxidizer to Fuel (Ratio)

Outer Gimbal

Oxygen Supply & Cabin Pressurization

Operations

Optical Tracking System

Outlet

(XVI) 5-l

OXID

0/F

P

PA

Pc

PCM

PCMTEA

PDS

PGNCS

HPPH

PIPA

PLSS

POSIT

PRESS

PRF

PROP

FWR

R

RCS

RCVR

RES

RF

RFI

RFSTF

RG

RGA

RNDZ

RR

RSS

RTG

S/C

SEL

SHe

SIN

SLA

Oxidizer

Oxidizer to Fuel

Pressure

Power Amplifier

Thrust-Chamber Pressure

Pulse Code Modulation

Pulse Code Modulation Timing Electronics Assembly

Power Distribution Section

Primary Guidance Navigation and Control Section

Hydrogen Potential

Phase

Pulsed Integrating Pendulous Accelerometer

Portable Life Support System

Position

Pressure

Pulse Repetition Frequency

Propulsion

Power

Required

Reaction Control Subsystem

Receiver

Resistance

Radio Frequency

Radio Frequency Interference

RF System Test Facility

Rate Gyro

Rate Gyro Assembly

Rendezvous

Rendezvous Radar

Root Sum Square

Radioisotope Thermal Generator

Spacecraft

Select

Supercritical Helium..

Sine

Spacecraft - LEM Adapter

(xvi)5-5

s/oSOL

S/S

ST

SURF

SYNCH

S/V

TCA

TOLER

TRUW

USB

VAB

VDC

VLV

VPI

VSR

W

WMS

X

X Feed

XMTR

X Pond

Y

Z

A

AV

Shut-off

Solenoid

Support Stand

Steerable

Surface

Synchronization

Space Vehicle

Thrust Chamber Assembly

Tolerance

Trunnion

Unified S-Band

Vehicle Assembly Building

Volts Direct Current

Valve

Valve Position Indicator

Voice Storage Recorder

Warning Light

Water Management Section

X-Axis (Vertical)

Crossfeed

Transmitter

Transponder

Y-Axis (Lateral)

Z-Axis (Fore-aft)

Delta

Change of Velocity

Grumman A i rc ra f t Engineering Corporat ion • Bethpage, New York

439

^h-2.6-1 unified s-band station capabilities 2-53 2.6-2 summary of ground station coverage - polar...

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